Ply stacking for solid laminate stiffened panels.

Abstract

To provide a composite panel for a fuselage, a wing, a tail, an empennage, etc. of an aircraft, such as a solid laminate stringer that connects to a supporting framework, the composite panel being resistant to damage caused by stresses and strains placed during use.SOLUTION: A composite panel 100 includes a plurality of segments 104, 106, 108, 110, where each segment includes a plurality of reinforcement plies such as carbon fiber plies. Collimated fiber bands (e.g., carbon fibers) within each reinforcement ply are oriented in a single direction, which can be, for example, 0°, 45°, 90° and -45°. Each segment can include a stack of reinforcement plies, where the collimated fiber bands of each reinforcement ply are oriented in one of these directions.SELECTED DRAWING: Figure 1

Description

[Technical field]

[0001] The present teachings relate generally to structural components such as composite panels used to manufacture aircraft or other products, and more particularly to composite structural component assemblies and methods for forming composite structural component assemblies. [Background technology]

[0002] Aircraft fuselages, wings, tails, empennages, etc. may be manufactured using composite panels such as solid laminate stringers connected to supporting frames. The composite panels may include a skin and an interior. An interior laminate at least partially forms the stringer. The skin and interior laminate may each be formed from multiple carbon fiber layers or plies, called prepregs, bonded together by a thermoplastic adhesive or a thermoplastic matrix. The skin of the composite panel is exposed and provides the exterior surface of the aircraft in use. The composite panels must be resistant to damage from stresses and strains resulting, for example, from aerodynamic loads and temperature changes imposed on the composite panel in use. Such stresses and strains may cause the composite panel to bend, fold, twist, expand, contract, etc., in use, which may result in delamination or other damage to the carbon fiber plies over time.

[0003] It would be a welcome addition to the art to provide composite panels that are more resistant to damage resulting from the stresses and strains experienced during use. Summary of the Invention

[0004] The following presents a simplified summary in order to provide a basic understanding of some aspects of one or more implementations of the present teachings. This summary is not an extensive overview, and is not intended to identify key or critical elements of the present teachings or to delineate the scope of the disclosure. Rather, its primary purpose is merely to present one or more concepts in a simplified form as a prelude to the more detailed description presented later.

[0005] In one implementation, a composite panel includes a composite panel segment. The composite panel segment may include a plurality of carbon fiber plies arranged along a length and a width of the composite panel segment. The length of the composite panel segment has a longitudinal axis. The plurality of carbon fiber plies are stacked into a carbon fiber stack that provides at least a portion of a thickness of the composite panel segment, the stack including at least eight carbon fiber plies. Each carbon fiber ply of the carbon fiber stack has a stiffness, the stiffness of each carbon fiber ply being determined by an orientation of the carbon fibers in each carbon fiber ply relative to the longitudinal axis. The plurality of carbon fiber plies of the carbon fiber stack have a plurality of different orientations relative to the longitudinal axis, whereby the plurality of carbon fiber plies of the carbon fiber stack may include a plurality of different stiffnesses. The various orientations of the plurality of carbon fiber plies of the carbon fiber stack are self-symmetrical or generally self-symmetrical about a midpoint of the carbon fiber stack across the thickness of the composite panel segment, the midpoint being parallel to the longitudinal axis, and the various orientations vary from self-symmetrical at no more than two carbon fiber plies when generally self-symmetrical.

[0006] Optionally, the longitudinal axis of the composite panel segment may be oriented at 0 degrees, and the various orientations of the carbon fiber plies include 0 degrees, 45 degrees, 90 degrees, and -45 degrees relative to the longitudinal axis of the composite panel segment. The various orientations of the carbon fiber plies of the carbon fiber laminate may be self-symmetrical across a thickness of the composite panel segment. Additionally, the composite panel segment may be a middle segment positioned between the base segment and the top segment, and 46% to 58% of the carbon fiber plies may be oriented at 0 degrees. The middle segment may have an axial modulus of elasticity of about 11.4 megapounds per square inch (msi) to about 13.4 msi. Additionally, the middle segment may have a Poisson's ratio of about 0.38 to about 0.50.

[0007] In one implementation, the plurality of carbon fiber plies of the carbon fiber laminate may include a plurality of repeatable groupings of carbon fiber plies, with each repeatable grouping of carbon fiber plies being repeated two, four, or six times. The various orientations of the plurality of carbon fiber plies of the carbon fiber laminate may be generally self-symmetric across the thickness of the composite panel segment. The carbon fiber laminate may include at least 16 carbon fiber plies.

[0008] In one implementation, the composite panel segment can be a mid-segment and the carbon fiber laminate can be a mid-segment carbon fiber laminate, and the composite panel can further include a base segment and a top segment, the mid-segment being positioned between the base segment and the top segment. In this implementation, the base segment can include a plurality of carbon fiber plies disposed along a length and width of the base segment, the length of the base segment being parallel to the longitudinal axis, and the plurality of carbon fiber plies of the base segment are stacked into a base segment carbon fiber laminate that provides at least a portion of a thickness of the base segment, and the base segment carbon fiber laminate can include at least 10 carbon fiber plies. Further, the carbon fiber plies of the base segment carbon fiber laminate may have a plurality of different orientations relative to the longitudinal axis, whereby the carbon fiber plies of the base segment carbon fiber laminate may include a plurality of different stiffnesses, where the various orientations of the carbon fiber plies of the base segment carbon fiber laminate are self-symmetric or generally self-symmetric about a midpoint of the base segment carbon fiber laminate across a thickness of the base segment, where the midpoint of the base segment is parallel to the longitudinal axis, where the various orientations of the carbon fiber plies of the base segment carbon fiber laminate vary from self-symmetric at no more than two carbon fiber plies when generally self-symmetric. Optionally, the longitudinal axis of the composite panel segment is oriented at 0 degrees, and the various orientations of the carbon fiber plies of the mid segment carbon fiber laminate and the carbon fiber plies of the base segment carbon fiber laminate include 0 degrees, 45 degrees, 90 degrees, and -45 degrees relative to the longitudinal axis of the composite panel segment. The middle segment may have an axial modulus of about 11.4 megapounds per square inch (msi) to about 13.4 msi, and the base and top segments may each have an axial modulus of about 10.0 msi to about 11.4 msi.

[0009] In one implementation, the middle segment may have a Poisson's ratio of about 0.38 to about 0.50, and the base segment and the top segment may each have a Poisson's ratio of about 0.36 to about 0.50. The top segment may include a plurality of carbon fiber plies arranged along the length and width of the top segment, and the length of the top segment may be parallel to the longitudinal axis. The plurality of carbon fiber plies of the top segment may be stacked into a top segment carbon fiber laminate that provides at least a portion of the thickness of the top segment. Further, the top segment carbon fiber laminate may include at least 10 carbon fiber plies, and the plurality of carbon fiber plies of the top segment carbon fiber laminate may have a plurality of different orientations with respect to the longitudinal axis, whereby the plurality of carbon fiber plies of the top segment carbon fiber laminate may include a plurality of different stiffnesses. The various orientations of the plurality of carbon fiber plies of the top segment carbon fiber laminate may be self-symmetrical or generally self-symmetrical around a midpoint of the top segment carbon fiber laminate across the thickness of the top segment, and the midpoint of the top segment carbon fiber laminate may be parallel to the longitudinal axis. The various orientations of the multiple carbon fiber plies of the upper segment carbon fiber laminate vary from being self-symmetrical at no more than two carbon fiber plies to being generally self-symmetrical, and the multiple carbon fiber plies of the upper segment carbon fiber laminate may be mirror symmetrical to the multiple carbon fiber plies of the base segment carbon fiber laminate.

[0010] In an optional implementation, the upper segment carbon fiber laminate, the intermediate segment carbon fiber laminate, and the base segment carbon fiber laminate can be generally self-symmetric around the midpoint of the intermediate segment carbon fiber laminate. Further, the upper segment carbon fiber laminate, the intermediate segment carbon fiber laminate, and the base segment carbon fiber laminate can form at least a part of the stringer of the composite panel. The composite panel can further include an outer panel segment, and the upper segment, the intermediate segment, and the base segment can be positioned on the side of the outer panel segment, and the base segment can be positioned between the intermediate segment and the outer panel segment. Further, the outer panel segment can include a plurality of carbon fiber plies arranged along the length and width of the outer panel segment, and the length of the outer panel segment can be parallel to the longitudinal axis. The plurality of carbon fiber plies of the outer panel segment can be laminated to form an outer panel segment carbon fiber laminate that provides at least a part of the thickness of the outer panel segment. Further, the outer panel segment carbon fiber laminate can include at least eight carbon fiber plies, and the plurality of carbon fiber plies of the outer panel segment carbon fiber laminate can have a plurality of different orientations with respect to the longitudinal axis, whereby the plurality of carbon fiber plies of the outer panel segment carbon fiber laminate can include a plurality of different rigidities. The various orientations of the plurality of carbon fiber plies of the outer panel segment carbon fiber laminate can be self-symmetric across the thickness of the outer panel segment around the midpoint of the outer panel segment carbon fiber laminate, and the midpoint of the outer panel segment carbon fiber laminate is parallel to the longitudinal axis.

[0011] In one implementation, the longitudinal axis of the composite panel segment is oriented at 0 degrees, and various orientations of the carbon fiber plies of the mid segment carbon fiber laminate, the base segment carbon fiber laminate, and the skin segment carbon fiber laminate include 0 degrees, 45 degrees, 90 degrees, and -45 degrees relative to the longitudinal axis of the composite panel segment, and about 46% to about 58% of the carbon fiber plies of the mid segment carbon fiber laminate may be oriented at 0 degrees. Optionally, about 40% to about 46% of the carbon fiber plies of each of the top segment carbon fiber laminate and the base segment carbon fiber laminate may be oriented at 0 degrees. Further optionally, the middle segment may have an axial modulus of about 11.4 megapounds per square inch (msi) to about 13.4 msi, the base segment and the top segment may each have an axial modulus of about 10.0 msi to about 11.4 msi, and the skin segment may have an axial modulus of about 7.7 msi to about 10.0 msi.

[0012] Optionally, the middle segment may have a Poisson's ratio of about 0.38 to about 0.50, the base segment and the top segment may each have a Poisson's ratio of about 0.36 to about 0.50, and the skin segment may have a Poisson's ratio of about 0.32 to about 0.48. The top segment, middle segment, and base segment may form at least a portion of a solid laminate, and the Poisson's ratio mismatch between the solid laminate and the skin segment may be in the range of about -0.06 to about 0.06.

[0013] In another implementation, a composite panel includes a composite panel segment having a plurality of reinforcing plies disposed along a length and width of the segment. The length of the composite panel segment has a longitudinal axis. The plurality of reinforcing plies are stacked into a reinforcing ply stack that spans at least a portion of a thickness of the composite panel segment. The plurality of reinforcing plies of the reinforcing ply stack include a plurality of different orientations relative to the longitudinal axis, and the different orientations of the plurality of reinforcing plies of the reinforcing ply stack are self-symmetric or generally self-symmetric about a midpoint of the reinforcing ply stack across the thickness of the composite panel segment, the midpoint being parallel to the longitudinal axis, and the different orientations vary from being self-symmetric at no more than two carbon fiber plies when being generally self-symmetric. Optionally, the plurality of reinforcing plies includes at least 16 reinforcing plies, and the orientations of the plurality of reinforcing plies are self-symmetric.

[0014] In another implementation, a method for forming a composite panel includes laminating a plurality of carbon fiber plies together to form a carbon fiber laminate of a composite panel segment, the carbon fiber laminate including at least 10 carbon fiber plies, and bonding the plurality of carbon fiber plies together. In this implementation, the plurality of carbon fiber plies are disposed along a length and a width of the composite panel segment, the plurality of carbon fiber plies forming at least a portion of a thickness of the composite panel segment, the length of the composite panel segment including a longitudinal axis. Further, each carbon fiber ply of the carbon panel laminate has a stiffness, the stiffness of each carbon fiber ply being determined by an orientation of the carbon fibers in each carbon fiber ply relative to the longitudinal axis, the plurality of carbon fiber plies of the carbon fiber laminate having a plurality of different orientations relative to the longitudinal axis, whereby the plurality of carbon fiber plies of the carbon fiber laminate include a plurality of different stiffnesses. Further, the various orientations of the plurality of carbon fiber plies of the carbon fiber laminate are self-symmetric or generally self-symmetric across the thickness of the composite panel segment about a midpoint of the carbon fiber laminate, the midpoint being parallel to the longitudinal axis. The various orientations vary from being self-symmetric at no more than two carbon fiber plies when generally self-symmetric. The method may optionally include orienting the longitudinal axis of the composite panel segment at 0 degrees, orienting each carbon fiber ply of the plurality of carbon fiber plies to have an orientation at one of 0 degrees, 45 degrees, 90 degrees, and -45 degrees relative to the longitudinal axis, and stacking at least 10 carbon fiber plies such that the various orientations of the plurality of carbon fiber plies of the carbon fiber laminate are self-symmetric.

[0015] The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate one implementation of the present teachings and, together with the description, serve to explain the principles of the disclosure. [Brief description of the drawings]

[0016] [Figure 1] FIG. 1 is a perspective view of a composite panel including a skin and a stringer in accordance with an exemplary implementation of the present teachings. [Diagram 2] FIG. 2 is a cross-sectional view taken along line 2-2 of FIG. [Diagram 3] FIG. 1 is a plan view of four parallelized fiber strips of four different reinforcing plies according to one implementation of the present teachings. [Figure 4] 1 is a table listing reinforcing plies and their fiber orientations in a reinforcing ply stack of a skin segment of a composite panel in accordance with an exemplary implementation of the present teachings. [Diagram 5] 1 is a table listing reinforcing plies and their fiber orientations in a mid-segment reinforcing ply stack that partially defines a stringer of a composite panel in accordance with an exemplary implementation of the present teachings. [Figure 6] 1 is a table listing reinforcing plies and their fiber orientations in a reinforcing ply stack of a base segment that partially provides a stringer of a composite panel, in accordance with an exemplary implementation of the present teachings. [Figure 7] 1 is a table listing reinforcing plies and their fiber orientations in a reinforcing ply stack of an upper segment that partially provides a stringer of a composite panel in accordance with an exemplary implementation of the present teachings. [Figure 8] 1 illustrates a runout for a stringer including a base segment and an upper segment in an exemplary implementation of the present teachings. [Figure 9] 1 is a table listing parameters and characteristics of various subcomponents of a composite panel according to an exemplary implementation of the present teachings. [Figure 10] 1 is a graph illustrating the axial modulus (i.e., stiffness) of various components of a composite panel in an exemplary implementation of the present teachings. [Figure 11] 1 is a schematic diagram of a stringer arrangement for a composite panel in an exemplary implementation of the present teachings. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0017] It should be noted that, rather than maintaining strict structural accuracy, detail, and scale, some details of the drawings have been simplified and illustrated to facilitate understanding of the present teachings.

[0018] Reference will now be made in detail to exemplary implementations of the present teachings, examples of which are illustrated in the accompanying drawings. Generally, and / or for convenience, the same reference numerals will be used throughout the drawings to refer to the same or like parts.

[0019] One implementation of the present teachings includes a composite panel and a method for manufacturing the composite panel. The composite panel can be a subcomponent of a larger structure (e.g., an aircraft, aerospace vehicle, or other structure). The composite panel can include multiple segments. Each segment includes multiple stacked plies of reinforcing material (i.e., reinforcing plies), which can be or include, for example, carbon fiber. In one implementation, the composite panel configuration is designed to be more resistant to delamination and thermal cracking, which can result from temperature mismatch, bending or warping of the composite panel during use, or other stresses and strains, than some conventional composite panels. The composite panel can include solid laminate stringers and configurations that facilitate runout of the stringers.

[0020] Figure 1 is a perspective view of a portion of a composite panel 100 according to one implementation of the present teachings, and Figure 2 is a cross-sectional view taken along line 2-2 in Figure 1. For purposes of illustration, the composite panel 100 includes a solid laminate stringer (hereinafter "stringer") 102, and a skin segment (i.e., skin laminate) 104, although other types of composite panels are contemplated that do not include the stringer 102 and / or the skin segment 104. Stringers and skins are well known, for example, in the field of aircraft manufacturing and other industries.

[0021] The stringer 102 in FIG. 1 includes a base segment 106, an intermediate segment 108, and an upper segment 110. As shown in FIGS. 1 and 2, the base segment 106 and the intermediate segment 108 are positioned between the outer plate segment 104 and the upper segment 110, and the intermediate segment 108 is positioned between the base segment 106 and the upper segment 110.

[0022] As described in more detail below, the outer plate segment 104, the base segment 106, the intermediate segment 108, and the upper segment 110 each include a plurality of reinforcing plies 200, 202, 204, and 206. For simplicity, only some of the reinforcing plies 200 - 206 that only partially extend across each segment 104 - 110 are shown, but it should be understood that each segment 104 - 110 includes at least eight reinforcing plies 200 - 206 that extend across most or all of each segment 104 - 110. In some structures, each segment 104 - 110, or one or more of the segments 104 - 110, may include at least 10, or at least 16, or at least 100, or at least 200, or at least 500, or at least 1000 reinforcing plies 200 - 206. Each of the plurality of reinforcing plies can be, for example, a plurality of carbon fiber plies. The plurality of reinforcing plies can be pre-impregnated with an adhesive resin (e.g., a thermoplastic resin) that bonds the plurality of reinforcing plies of each segment together and can be called a prepreg. Each ply of the plurality of reinforcing plies 200 - 208 extends across the length and width of each segment 104 - 110 and is oriented as described. The plurality of plies 200 - 208 are each laminated (i.e., stacked) along the thickness of each segment 104 - 110 to form a reinforcing ply laminate (e.g., a carbon fiber laminate). In the orientation of FIG. 2, each reinforcing ply extends horizontally and each reinforcing ply laminate includes a plurality of reinforcing plies stacked vertically.

[0023] Each segment 104-110 may be formed using an automated fiber placement (AFP) process. In an automated fiber placement (AFP) process, a computer-controlled fiber placement head of a fiber placement system deposits a plurality of prepreg tows that form parallelized fiber bands on a surface. In some processes, the parallelized fiber bands are deposited on a mandrel that provides or defines the final component to be produced. In these processes, the segments 104-110 are chemically bonded and / or mechanically fastened together. In other processes, the skin segment 104 may be deposited and manufactured first, then the base segment 106 is deposited on the skin segment 104, the middle segment 108 is deposited on the base segment 106, and the top segment 110 is deposited on the middle segment 108. Alternatively, the order may be reversed, with the top segment 110 deposited first and the skin segment 104 deposited last on the base segment 106.

[0024] 1 illustrates a longitudinal axis "A" that extends through and is parallel to the length of the composite panel 100 and the length of the stringers 102. For purposes of this disclosure, the longitudinal axis A that is through and parallel to the composite panel 100 is defined to be at 0 degrees (i.e., axial) relative to the composite panel 100. During deposition of each collimated fiber strip, the fiber placement head may move in a 0 degree direction and deposit the collimated fiber strip at 0 degrees such that the collimated fiber strip is oriented at 0 degrees. In this orientation, the fiber placement head needs to make a minimum number of passes across the composite panel to deposit multiple collimated fiber strips that form a single reinforcing ply.

[0025] FIG. 3 is a schematic plan view of a composite panel 100 with its length "L" or axis "A" oriented at 0 degrees from top to bottom and its width "W" oriented from left to right. During the deposition of each parallelized fiber strip 300-306, the direction of movement of the fiber placement head can be changed to deposit the parallelized fiber strips 300-306 in a variety of different orientations. FIG. 3 shows four parallelized fiber strips 300-306 deposited during four separate passes of the fiber placement head in four different directions. The parallelized fiber strips 300, 302, 304, and 306 are oriented at 0 degrees, 45 degrees, 90 degrees, and -45 degrees, respectively. For simplicity, FIG. 3 shows the results of only one pass of the fiber placement head for each reinforcing ply, but it should be understood that a sufficient number of passes of the fiber placement head are made to complete each reinforcing ply. Furthermore, the four collimated fiber bands 300-306 and their corresponding reinforcing plies are stacked vertically (ie, perpendicular to the plane of the page).

[0026] The particular orientation of the paralleled fiber strips 300-306 that are laid down to form one of the segments 104-110 determines the stiffness of the segment 104-110. For example, a segment including multiple reinforcing plies with paralleled fiber strips oriented at only 0 degrees will have a relatively stiff segment, while a segment including multiple reinforcing plies oriented at only 90 degrees will have a relatively flexible segment.

[0027] The present teachings include a composite panel having a plurality of segments, the stiffness of each segment being controlled during design and manufacture of the composite panel by selecting an orientation of the parallelized fiber bands within each of the plurality of reinforcing plies that make up the segment. Additionally, the orientation of the parallelized fiber bands of the reinforcing plies within each segment is designed and selected to improve adhesion between adjacent reinforcing plies as compared to some conventional designs, thereby improving overall adhesion of each segment and reducing delamination within each segment. Additionally, the stiffness of each of the plurality of segments that make up the composite panel is designed to improve adhesion between adjacent segments, thereby reducing stiffness mismatch between adjacent segments and reducing separation of the composite panel between adjacent segments.

[0028] In one implementation of the present teachings, the orientation of the reinforcing plies in each individual segment across the thickness of each segment is designed, controlled, and / or selected to be self-symmetrical or approximately self-symmetrical. For the purposes of this disclosure, unless otherwise specified, in a self-symmetrical segment, the entire reinforcing plies that make up the self-symmetrical segment are mirror symmetrical around the midpoint of the thickness of the self-symmetrical segment. In a generally self-symmetrical segment, no more than two reinforcing plies of the reinforcing plies that make up the generally self-symmetrical segment deviate from or differ from mirror symmetry around the midpoint of the thickness of the segment. In other words, in a generally self-symmetrical segment, changing the orientation of no more than two reinforcing plies (i.e., one ply or two plies) results in the segment being self-symmetrical.

[0029] Furthermore, the term "self-symmetric" when referring to a plurality of reinforcing plies forming the entire thickness of two or more segments indicates that the plurality of reinforcing plies forming the two or more segments are mirror symmetric about the midpoint of the two or more segments, whereas "mostly self-symmetric" indicates that in the plurality of reinforcing plies forming the two or more segments, no more than two of the plurality of reinforcing plies constituting the mostly self-symmetric stack deviate from or differ from the mirror symmetry about the midpoint of the thickness of the two or more segments. In other words, changing the orientation of no more than two reinforcing plies (i.e., one ply or two plies) across two or more segments results in the plurality of reinforcing plies forming the two or more segments being self-symmetric. The two or more segments that are self-symmetric or mostly self-symmetric may be positioned adjacent to each other or may have other configurations such that another segment is positioned in between.

[0030] Additionally, the "midpoint" of a segment (or two or more segments) is the midpoint relative to the number of reinforcing plies that make up the reinforcing ply stack. In a segment (or two or more segments) that includes an odd number of reinforcing plies, the midpoint is the middle reinforcing ply, and in a segment (or two or more segments) that includes an even number of reinforcing plies, the midpoint is between two reinforcing plies at the midpoint of the segment (or the midpoint of two or more segments), with an equal number of reinforcing plies on either side of the midpoint. Additionally, "corresponding" reinforcing plies refer to two reinforcing plies that are equidistant from and on either side of the common midpoint.

[0031] As mentioned above, in one implementation of the present teachings, all stiffening plies that make up two or more segments of a composite panel can be self-symmetrical or generally self-symmetrical across the thickness of the composite panel of the two or more segments. With respect to the example of Figures 1 and 2, the three segments 106-110 that make up the stringer 102 can be self-symmetrical or generally self-symmetrical. Furthermore, the stiffness of the stringer 102 (i.e., the segments 106-110 that make up the stringer 102) can be matched to the stiffness of the skin panel segment 104. Matching the stiffness of these two structures (i.e., the skin panel segment 104 and the stringer 102) can reduce or eliminate stiffness mismatches that can pull the stringer away from the skin panel.

[0032] 4-7 are tables listing exemplary reinforcement ply orientations for the four segments 104-110 of FIG. 1, with FIG. 4 showing an exemplary reinforcement ply orientation for the skin segment 104, FIG. 5 showing an exemplary reinforcement ply orientation for the mid-segment 108, FIG. 6 showing an exemplary reinforcement ply orientation for the base segment 106, and FIG. 7 showing an exemplary reinforcement ply orientation for the top segment 110. As shown in FIGS. 4-7, the skin segment 104 includes 40 reinforcement plies, the mid-segment 108 includes 32 reinforcement plies, and the base segment 106 and top segment 110 include 14 plies each. It should be understood that the segments 104-110 may each have more or less plies than those shown. For example, in one implementation, the segments 104-110 may each include at least 8 plies and may include more than 1000 plies. In each of these implementations, the reinforcement plies within each segment may be self-symmetrical or generally symmetrical. Additionally, the reinforcing plies across all segments making up the composite panel are self-symmetrical or generally symmetrical.

[0033] FIG. 4 is a table listing the reinforcing plies and their fiber orientations in a self-symmetric skin segment 104 having 40 reinforcing plies 200 (shown in FIG. 2). As shown in FIG. 4, the bottom ply (ply #1) is oriented in the same direction (45 degrees) as the top ply (ply #40), ply #2 is oriented in the same direction (90 degrees) as ply #39, and so on until a midpoint 400 between reinforcing plies #20 and #21, where reinforcing plies #20 and #21 are both oriented at 0 degrees. Additionally, skin segment 104 may include reinforcing plies 200 constructed from about 25% to about 40% collimated fiber bands oriented in the 0 degree direction. With these parameters, skin segment 104 will have an axial modulus of elasticity of about 7.7 megapounds per square inch (msi) to about 10.0 msi, with a Poisson's ratio of about 0.32 to about 0.48.

[0034] FIG. 5 is a table listing the reinforcing plies and their fiber orientations in a self-symmetrical middle segment 108 having 32 (shown in FIG. 2) reinforcing plies 204. As shown in FIG. 5, the bottom ply (ply #1) is oriented in the same direction (0 degrees) as the top ply (ply #32), ply #2 is oriented in the same direction (0 degrees) as ply #32, and so on until a midpoint 500 between reinforcing plies #16 and #17, where reinforcing plies #20 and #21 are both oriented at 0 degrees. The ply stack shown in FIG. 5 provides a repeatable group of reinforcing plies that can be replicated or repeated in any even multiple (2x, 4x, 6x, etc.) to create a thicker middle segment 108. This allows the entire solid laminate (e.g., stringer 102) adjacent to the skin segment 104 to have a thickness selected or designed based on expected loads, deformations, and damage protection in service. Additionally, the middle segment 108 may include a reinforcing ply 204 constructed from about 46% to about 58% 0 degree oriented collimated fiber bands. With these parameters, the middle segment will have an axial modulus of about 11.4 msi to about 13.4 msi, and a Poisson's ratio of about 0.38 to about 0.50.

[0035] FIG. 6 is a table listing the reinforcing plies and their fiber orientations in a generally self-symmetric base segment 106 having fourteen (shown in FIG. 2) reinforcing plies 202. As shown in FIG. 6, the bottom most ply (ply #1) is oriented in the same direction (45 degrees) as the top most ply (ply #14), ply #2 is oriented in the same direction (90 degrees) as ply #13, and so on up to corresponding reinforcing plies #6 (oriented at 45 degrees) and #9 (oriented at -45 degrees) which are not symmetric. The remaining reinforcing plies 7 and 8 are symmetric about the midpoint 600 between them. Thus, the base segment 106 (shown in FIG. 2) is generally self-symmetric because by changing one ply (i.e., changing ply 6 to -45 degrees or changing ply 9 to 45 degrees), the base segment 106 becomes self-symmetric. Additionally, the base segment 106 may include a reinforcement ply 202 constructed from about 40% to about 46% 0 degree oriented collimated fiber bands. With these parameters, the base segment 106 will have an axial modulus of elasticity of about 10.0 msi to about 11.4 msi, and a Poisson's ratio of about 0.36 to about 0.50.

[0036] FIG. 7 is a table listing the reinforcing plies and their fiber orientations in a generally self-symmetrical top segment 110 having 14 reinforcing plies 206 (shown in FIG. 2). As shown in FIG. 7, the bottom ply (ply #N-13) is oriented in the same direction (45 degrees) as the top ply (ply #N), ply #N-1 is oriented in the same direction (90 degrees) as ply #N-12, and so on up to corresponding reinforcing plies N-8 (oriented at -45 degrees) and N-5 (oriented at 45 degrees) which are not symmetric. The remaining reinforcing plies N-7 and N-6 are symmetric about the midpoint 700 between them. Thus, the top segment 110 (shown in FIG. 2) is generally self-symmetrical since by changing one ply (i.e., changing ply N-8 to 45 degrees or changing ply N-5 to -45 degrees), the top segment 110 becomes self-symmetrical. Additionally, the top segment 110 may include a reinforcing ply 206 constructed from about 40% to about 46% 0 degree oriented collimated fiber bands. With these parameters, the top segment 110 will have an axial modulus of about 10.0 msi to about 11.4 msi, and a Poisson's ratio of about 0.36 to about 0.50.

[0037] Additionally, segments 104-110 will include reinforcing plies constructed from paralleled fiber bands oriented at a transverse angle of about 12% or greater (ie, 90 degrees).

[0038] Forming the skin segments 104, base segments 106, middle segments 108, and top segments 110 with parallelized fiber bands in the proportions set forth above results in the axial modulus and Poisson's ratios set forth above. Forming the segments to have the axial modulus and Poisson's ratios in the particular ranges set forth above results in an integrated structure that has increased resistance to delamination, reduced global waviness, reduced weight, and reduced manufacturing and material costs compared to conventional methods using conventional techniques. Forming the segments 104-110 with a lower axial modulus would result in a structure with increased weight, manufacturing time, and cost. Forming the segments 104-110 with a higher axial modulus would result in a structure with increased propensity for delamination due to increased stress mismatch between the skin segments 104 and the stringers 102. Forming the segments 104-110 with a lower Poisson's ratio would result in a structure with reduced resistance to delamination resulting from Poisson's ratio mismatch between the structures. Forming segments 104-110 with a higher Poisson's ratio than described above results in an increased mismatch of Poisson's ratios, reduced resistance to delamination, and increased overall waviness.

[0039] 6 and 7, the base segment 106 and the top segment 110 are self-symmetrical (i.e., mirror symmetrical) with respect to one another about the midpoint 650 therebetween. At a solid laminate termination region or runout, such as runout 800 (e.g., at the end of the stringer 102) shown in FIG. 8, the base segment 106 and the top segment 110 are joined to form the self-symmetric runout 800. For example, during formation of the stringer 102, the multiple plies 204 forming the middle segment 108 may be formed at different lengths such that the middle segment 108 is progressively thinner such that at the end of the stringer 102, the runout 800 includes only the base segment 106 and the top segment 110. In other words, the plies that make up the self-symmetrical reinforcement ply stack of the middle segment 108 are sequentially dropped into corresponding pairs of reinforcement plies (e.g., plies 1 and 32, then plies 2 and 31, etc.) such that only the base segment 106 and the top segment 110 remain to form the runout 800. As shown in FIG. 8, the base segment 106 and the top segment 110 have the same thickness. Furthermore, as discussed above, the base segment 106 is a mirror image of the top segment 110 with respect to a midpoint 802 of the interface between the two segments 106, 110. It should be understood that the midpoint 802 is similar to the midpoint 650 of FIG. 6.

[0040] Furthermore, because the intermediate segment 108 is self-symmetrical about the midpoint 500, the stringer 102 formed from the base segment 106, the intermediate segment 108, and the upper segment 110 is self-symmetrical when the intermediate segment 108 is positioned between the base segment 106 and the upper segment 110.

[0041] FIG. 9 is a table illustrating the axial fiber % (i.e., percentage of paralleled fiber bands oriented at 0 degrees), transverse fiber % (i.e., percentage of paralleled fiber bands oriented at 90 degrees), axial modulus, Poisson's ratio, and Poisson's ratio mismatch for each segment 104-110 for the skin in an exemplary implementation described herein. The Poisson's ratio mismatch between the skin segment 104 and the base segment 106 is within a range of about -0.06 to about 0.06. Additionally, the Poisson's ratio mismatch between the base segment 106 and the middle segment, and between the top segment 110 and the middle segment 108 is within a range of about -0.06 to about 0.06. This relatively low Poisson's ratio mismatch reduces or prevents the stringer 102 from delaminating from the skin segment 104, and further reduces or prevents the base segment 106 and the top segment 110 from delaminating from the middle segment 108.

[0042] In the composite panel 100 of Figure 1, the overall stiffness of the composite panel varies progressively from the relatively soft skin segments 104 to the relatively stiffer base segments 106 and top segments 110 to the stiffer middle segment 108, which are stiffer than the skin segments 104 but not as stiff as the middle segment 108. The stiffness of each segment 104-110 in an exemplary implementation is shown in Figure 10, illustrating the gradual change in axial modulus between the segments. This gradation reduces the stiffness mismatch between adjacent segments and improves resistance to delamination, thereby reducing or eliminating separation and delamination between the segments.

[0043] The proposed overall combination and sequence of stiffening ply stacks from the skin segment 104 to the segments 106-110 forming the stringer 102 may reduce the overall waviness of the stiffened composite panel 100 and improve the structural quality of the composite panel. In one implementation, the overall waviness of the composite panel 100 may be reduced by 30% or more, or 40% or more, compared to composite panels manufactured using conventional techniques. This reduction in the overall waviness of the composite panel may result in reduced assembly operations and rework. Additionally, reduced overall waviness reduces routine costs and improves workforce safety, for example, by reducing industrial accidents resulting from working in restricted working spaces during assembly. Reduced overall waviness reduces or eliminates the need to fill reduced gaps between parts or structures caused by overall waviness.

[0044] As discussed above, the collimated fiber bands 300-306 of FIG. 3 are deposited using a fiber placement head of a fiber placement system. The fiber placement head makes multiple passes over the mandrel to deposit each reinforcing ply. Once the first ply is completed, the fiber placement head may deposit a second ply over the first ply. Reinforcing plies with a 0 degree orientation require the least amount of time to deposit, since the 0 degree orientation requires the fiber placement head to make the fewest number of passes over the mandrel to complete the reinforcing ply. Conversely, reinforcing plies with a 90 degree orientation require the most amount of time to deposit, since the 90 degree orientation requires the fiber placement head to make the greatest number of passes over the mandrel to complete the reinforcing ply.

[0045] In one implementation of the present teachings, a common multi-ply component or group of reinforcing plies may be preassembled (i.e., pre-manufactured) and stored for later use. The various needed multi-ply components can be removed from storage and assembled into the final component.

[0046] For example, FIG. 11 shows a plurality of multi-ply groups 1100. Each multi-ply group is designated by a letter A through J (there are 10 different multi-ply groups). FIG. 11 further shows a stringer 102 including a base segment 106, a middle segment 108, and a top segment 110. Each of the plurality of multi-ply groups A-J may be pre-assembled and stored until the composite panel 100 including the stringer 102 is to be manufactured. A required quantity of each multi-ply group A-J is ordered and delivered to an assembly site prior to building the composite panel 100. Build begins by placing the first required multi-ply group of the stringer 102 on a mandrel 1102. The first required multi-ply group is the three ply groups A of the base segment 106 in this exemplary implementation. The next required reinforcement ply is oriented at 0 degrees. This can be deposited relatively quickly as described above. Thus, the reinforcement ply oriented at 0 is deposited using the fiber placement head 1104 of the fiber placement system. Next, as shown in Figure 11, each required multi-ply group A-J is placed and each required reinforcing ply oriented at 0 degrees is deposited to form the base segment 106, the middle segment 108, and the top segment 110, respectively. In the exemplary implementation of Figure 11, after the last multi-ply group B is placed, the formation of this portion of the composite panel 100 is complete. Additional steps may continue to be completed as required by the particular process (e.g., prepreg debulking or resin curing).

[0047] Prefabricating the multi-ply groups reduces the manufacturing time required on the mandrel, thus increasing production throughput. Deposition of a 0 degree collimated fiber band requires a minimal number of passes of the fiber placement head 1104 over the mandrel 1102 and requires minimal time compared to forming collimated fiber bands of 45 degree, 90 degree, and 45 degree orientations. However, one or more multi-ply groups including one or more 0 degree reinforcing plies are contemplated and illustrated, for example, in multi-ply groups C-F. Additionally, while FIG. 11 illustrates ten multi-ply groups A-J having two or three reinforcing plies each, it should be understood that any number of multi-ply groups having any number of reinforcing plies in each group are contemplated.

[0048] Thus, in a method for forming a composite panel, a plurality of reinforcing plies (e.g., carbon fiber plies including parallelized fiber strips including carbon fibers) can be stacked together to form a carbon fiber laminate of a composite panel segment. The carbon fiber laminate may include, for example, at least 8 carbon fiber plies, or at least 10 carbon fiber plies, or at least 16 carbon fiber plies, or at least 100 carbon fiber plies, or at least 200 carbon fiber plies, or at least 1000 carbon fiber plies. For example, the carbon fiber plies are pre-impregnated, the carbon fiber plies are bonded together using a process including debulking the carbon fiber plies and curing the resin. The carbon fiber plies are disposed along the length and width of the composite panel segment and form at least a portion of the thickness of the composite panel segment. The length of the composite panel segment includes a longitudinal axis, and each carbon fiber ply of the carbon panel laminate includes a stiffness determined by the orientation of the carbon fibers in each carbon fiber ply relative to the longitudinal axis. The carbon fiber plies of the carbon fiber laminate have a plurality of different orientations relative to the longitudinal axis, whereby the carbon fiber plies of the carbon fiber laminate include a plurality of different stiffnesses. The various orientations of the carbon fiber plies of the carbon fiber laminate are self-symmetrical or generally self-symmetrical across the thickness of the composite panel segment about a midpoint of the carbon fiber laminate, the midpoint being parallel to the longitudinal axis. The various orientations vary from self-symmetrical in no more than two carbon fiber plies to generally self-symmetrical.

[0049] Furthermore, the present disclosure includes embodiments according to the following clauses:

[0050] Clause 1 A composite panel (100), comprising: a composite panel segment (104-110) comprising a plurality of carbon fiber plies (200-206) disposed along a length and a width of the composite panel segment (104-110), the length of the composite panel segment (104-110) comprising a longitudinal axis (A); the plurality of carbon fiber plies (200-206) are stacked into a carbon fiber laminate that provides at least a portion of the thickness of the composite panel segment (104-110); the carbon fiber laminate comprises at least eight carbon fiber plies (200-206); each carbon fiber ply of the carbon fiber laminate has a stiffness, the stiffness of each carbon fiber ply being determined by the orientation of the carbon fibers relative to the longitudinal axis (A) within each carbon fiber ply; the plurality of carbon fiber plies (200-206) of the carbon fiber laminate have a plurality of different orientations relative to the longitudinal axis (A), whereby the plurality of carbon fiber plies (200-206) of the carbon fiber laminate have a plurality of different stiffnesses; A composite panel (100) wherein the various orientations of the plurality of carbon fiber plies (200-206) of the carbon fiber laminate are self-symmetrical or generally self-symmetrical around a midpoint (400, 500, 600) of the carbon fiber laminate across a thickness of the composite panel segment (104-110), the midpoint (400, 500, 600) being parallel to the longitudinal axis (A), and the various orientations vary from being self-symmetrical for no more than two carbon fiber plies (200-206) when generally self-symmetrical.

[0051] Clause 2 the longitudinal axes (A) of the composite panel segments (104-110) are oriented at 0 degrees; 2. The composite panel (100) of claim 1, wherein the various orientations of the plurality of carbon fiber plies (200-206) include 0 degrees, 45 degrees, 90 degrees, and -45 degrees relative to the longitudinal axis (A) of the composite panel segment (104-110).

[0052] Clause 3 3. The composite panel (100) of claim 2, wherein the various orientations of the plurality of carbon fiber plies (200-206) of the carbon fiber laminate are self-symmetrical across a thickness of the composite panel segment (104-110).

[0053] Clause 4 The composite panel segments (104-110) are intermediate segments (108) positioned between a base segment (106) and an upper segment (110), the intermediate segment (108) having an axial modulus of elasticity of about 11.4 megapounds per square inch (msi) to about 13.4 msi and a Poisson's ratio of about 0.38 to about 0.50; 4. The composite panel (100) of claim 2 or 3, wherein between 46% and 58% of the plurality of carbon fiber plies (200-206) are oriented at 0 degrees.

[0054] Clause 5 5. The composite panel (100) of any one of clauses 1 to 4, wherein the plurality of carbon fiber plies (200-206) of the carbon fiber laminate comprises a plurality of repeatable groupings of carbon fiber plies (200-206), each repeatable grouping of carbon fiber plies (200-206) being repeated 2, 4, or 6 times.

[0055] Clause 6 3. The composite panel (100) of claim 1 or 2, wherein the various orientations of the plurality of carbon fiber plies (200-206) of the carbon fiber laminate are generally self-symmetrical across a thickness of the composite panel segment (104-110).

[0056] Clause 7 the composite panel segment (104-110) is a mid-segment (108), the carbon fiber laminate is a mid-segment carbon fiber laminate, and the composite panel (100) is a base segment (106) and an upper segment (110), the intermediate segment (108) being positioned between the base segment (106) and the upper segment (110); The base segment (106) a plurality of carbon fiber plies (202) disposed along a length and width of the base segment (106), the length of the base segment (106) being parallel to the longitudinal axis (A); the plurality of carbon fiber plies (202) of the base segment (106) are stacked into a base segment carbon fiber laminate that provides at least a portion of the thickness of the base segment (106); the base segment carbon fiber laminate comprises at least ten carbon fiber plies (202); the plurality of carbon fiber plies (202) of the base segment carbon fiber laminate have a plurality of different orientations relative to the longitudinal axis (A), whereby the plurality of carbon fiber plies (202) of the base segment carbon fiber laminate have a plurality of different stiffnesses; 4. The composite panel (100) of any one of clauses 1 to 3, wherein the various orientations of the plurality of carbon fiber plies (200-206) of the base segment carbon fiber laminate are self-symmetrical or generally self-symmetrical across a thickness of the base segment (106) about a midpoint (600) of the base segment carbon fiber laminate, the midpoint (600) of the base segment being parallel to the longitudinal axis (A), and the various orientations of the plurality of carbon fiber plies (200-206) of the base segment carbon fiber laminate vary from being self-symmetrical for no more than two carbon fiber plies (200-206) when generally self-symmetrical.

[0057] Clause 8 the longitudinal axes (A) of the composite panel segments (104-110) are oriented at 0 degrees; 8. The composite panel (100) of claim 7, wherein the various orientations of the plurality of carbon fiber plies (204) of the intermediate segment carbon fiber laminate and the plurality of carbon fiber plies (202) of the base segment carbon fiber laminate include 0 degrees, 45 degrees, 90 degrees, and -45 degrees relative to the longitudinal axis (A) of the composite panel segment (104-110).

[0058] Clause 9 the intermediate segment (108) has an axial modulus of elasticity of about 11.4 megapounds per square inch (msi) to about 13.4 msi; the base segment (106) and the top segment (110) each have an axial modulus of about 10.0 msi to about 11.4 msi; 9. A composite panel (100) according to clause 7 or 8.

[0059] Clause 10 the intermediate segment (108) has a Poisson's ratio of about 0.38 to about 0.50; the base segment (106) and the top segment (110) each have a Poisson's ratio of about 0.36 to about 0.50; 10. The composite panel (100) of any one of clauses 7 to 9.

[0060] Clause 11 The upper segment (110) a plurality of carbon fiber plies (206) disposed along a length and width of the upper segment (110), the length of the upper segment (110) being parallel to the longitudinal axis (A); the plurality of carbon fiber plies (206) of the upper segment (110) are stacked into an upper segment carbon fiber stack that provides at least a portion of the thickness of the upper segment (110); the upper segment carbon fiber laminate comprising at least ten carbon fiber plies (206); the plurality of carbon fiber plies (206) of the upper segment carbon fiber laminate have a plurality of different orientations relative to the longitudinal axis (A), whereby the plurality of carbon fiber plies (206) of the upper segment carbon fiber laminate have a plurality of different stiffnesses; the various orientations of the plurality of carbon fiber plies (206) of the upper segment carbon fiber laminate are self-symmetrical or generally self-symmetrical around a midpoint (700) of the upper segment carbon fiber laminate across a thickness of the upper segment (110), the midpoint (700) of the upper segment carbon fiber laminate being parallel to the longitudinal axis (A), and the various orientations of the plurality of carbon fiber plies (206) of the upper segment carbon fiber laminate vary from being self-symmetrical in any two or less carbon fiber plies (206) when generally self-symmetrical; 11. The composite panel (100) of any one of clauses 7 to 10, wherein the plurality of carbon fiber plies (206) of the upper segment carbon fiber laminate are mirror images of the plurality of carbon fiber plies (202) of the base segment carbon fiber laminate.

[0061] Clause 12 the top segment carbon fiber laminate, the mid-segment carbon fiber laminate, and the base segment carbon fiber laminate are generally self-symmetrical about the midpoint (500) of the mid-segment carbon fiber laminate; The composite panel (100) of any one of clauses 7 to 11, wherein the upper segment carbon fiber laminate, the mid segment carbon fiber laminate, and the base segment carbon fiber laminate form at least a portion of a stringer (102) of the composite panel (100).

[0062] Clause 13 a skin segment (104), wherein the upper segment (110), the intermediate segment (108), and the base segment (106) are positioned on sides of the skin segment (104), and the base segment (106) is positioned between the intermediate segment (108) and the skin segment (104); The skin segment (104) a plurality of carbon fiber plies (200) disposed along a length and width of the skin segment (104), the length of the skin segment (104) being parallel to the longitudinal axis (A); the plurality of carbon fiber plies (200) of the skin segment (104) are stacked into a skin segment carbon fiber laminate that provides at least a portion of the thickness of the skin segment (104); the skin segment carbon fiber laminate comprises at least eight carbon fiber plies (200); the plurality of carbon fiber plies (200) of the skin segment carbon fiber laminate have a plurality of different orientations relative to the longitudinal axis (A), whereby the plurality of carbon fiber plies (200) of the skin segment carbon fiber laminate have a plurality of different stiffnesses; 13. The composite panel (100) of any one of claims 7 to 12, wherein the various orientations of the plurality of carbon fiber plies (200) of the skin segment carbon fiber laminate are self-symmetrical around a midpoint (400) of the skin segment carbon fiber laminate through a thickness of the skin segment (104), the midpoint (400) of the skin segment carbon fiber laminate being parallel to the longitudinal axis (A).

[0063] Clause 14 the longitudinal axes (A) of the composite panel segments (104-110) are oriented at 0 degrees; the various orientations of the plurality of carbon fiber plies (204) of the intermediate segment carbon fiber laminate, the plurality of carbon fiber plies (202) of the base segment carbon fiber laminate, and the plurality of carbon fiber plies (200) of the skin segment carbon fiber laminate include 0 degrees, 45 degrees, 90 degrees, and -45 degrees relative to the longitudinal axis (A) of the composite panel segment (104-110); About 46% to about 58% of the plurality of carbon fiber plies (204) of the mid-segment carbon fiber laminate are oriented at 0 degrees; 14. The composite panel (100) of claim 13, wherein about 40% to about 46% of the plurality of carbon fiber plies (206, 202) of each of the upper segment carbon fiber laminate and the base segment carbon fiber laminate are oriented at 0 degrees.

[0064] Clause 15 the intermediate segment (108) having an axial modulus of elasticity of about 11.4 megapounds per square inch (msi) to about 13.4 msi and a Poisson's ratio of about 0.38 to about 0.50; the base segment (106) and the top segment (110) each have an axial modulus of elasticity of about 10.0 msi to about 11.4 msi; the base segment (106) and the top segment (110) each have a Poisson's ratio of about 0.36 to about 0.50; 15. The composite panel (100) of claim 13 or 14, wherein the skin segment (104) has an axial modulus of about 7.7 msi to about 10.0 msi.

[0065] Clause 16 the top segment (110), the middle segment (108), and the base segment (106) form at least a portion of a solid laminate; 16. The composite panel (100) of any one of clauses 13 to 15, wherein a Poisson's ratio mismatch between the solid laminate and the skin segments (104) is within a range of about -0.06 to about 0.06.

[0066] Clause 17 A composite panel (100), comprising: a composite panel segment (104-110) comprising a plurality of reinforcing plies (200-206) disposed along a length and a width of the segment (104-110), the length of the composite panel segment (104-110) comprising a longitudinal axis (A); the plurality of reinforcing plies (200-206) are stacked into a reinforcing ply stack that provides at least a portion of the thickness of the composite panel segment (104-110); the plurality of reinforcing plies (200-206) of the reinforcing ply stack have a plurality of different orientations relative to the longitudinal axis (A); A composite panel (100) wherein the various orientations of the plurality of reinforcing plies (200-206) of the reinforcing ply stack are self-symmetrical or generally self-symmetrical around a midpoint (400, 500, 600, 700) of the reinforcing ply stack across a thickness of the composite panel segment (104-110), the midpoint (400, 500, 600, 700) being parallel to the longitudinal axis (A), and the various orientations vary from being self-symmetrical in no more than two carbon fiber plies (200-206) when generally self-symmetrical.

[0067] Clause 18 said plurality of reinforcing plies (200-206) comprising at least 16 reinforcing plies (200-206); 18. The composite panel (100) of claim 17, wherein the various orientations of the plurality of reinforcing plies (200-206) are self-symmetrical.

[0068] Clause 19 A method for forming a composite panel (100), comprising: laminating together a plurality of carbon fiber plies (200-206) to form a carbon fiber stack of a composite panel segment (104-110), said carbon fiber stack comprising at least ten carbon fiber plies (200-206); Bonding a plurality of said carbon fiber plies (200-206) together. Including, the plurality of carbon fiber plies (200-206) are disposed along the length and width of the composite panel segment (104-110); the plurality of carbon fiber plies (200-206) forming at least a portion of the thickness of the composite panel segment (104-110); the length of the composite panel segment (104-110) includes a longitudinal axis (A); each carbon fiber ply of the carbon panel laminate has stiffness; the stiffness of each carbon fiber ply is determined by the orientation of the carbon fibers in each carbon fiber ply relative to the longitudinal axis (A); the plurality of carbon fiber plies (200-206) of the carbon fiber laminate have a plurality of different orientations relative to the longitudinal axis (A), whereby the plurality of carbon fiber plies (200-206) of the carbon fiber laminate have a plurality of different stiffnesses; the various orientations of the plurality of carbon fiber plies (200-206) of the carbon fiber laminate are self-symmetrical or generally self-symmetrical around a midpoint (400, 500, 600, 700) of the carbon fiber laminate through a thickness of the composite panel segment (104-110); The method of claim 1, wherein said intermediate points (400, 500, 600, 700) are parallel to said longitudinal axis (A) and said various orientations vary from being self-symmetrical in no more than two carbon fiber plies (200-206) when said intermediate points (400, 500, 600, 700) are parallel to said longitudinal axis (A) and said various orientations are generally self-symmetrical.

[0069] Clause 20 orienting the longitudinal axis (A) of the composite panel segments (104-110) at 0 degrees; orienting each carbon fiber ply of the plurality of carbon fiber plies (200-206) to have an orientation at one of 0 degrees, 45 degrees, 90 degrees, and -45 degrees relative to the longitudinal axis (A); stacking at least ten carbon fiber plies (200-206) such that the various orientations of the plurality of carbon fiber plies (200-206) of the carbon fiber laminate are self-symmetrical; 20. The method of claim 19, further comprising:

[0070] Although the numerical ranges and parameters setting forth the broad scope of the present teachings are approximate, the numerical values ​​set forth in the specific examples are set forth as precisely as possible. However, any numerical value inherently contains certain errors necessarily resulting from the standard deviation found in each testing measurement. Moreover, it should be understood that all ranges disclosed herein include any and all subranges contained therein. For example, a range "less than 10" can include any and all subranges between (and including) a minimum value of 0 and a maximum value of 10, that is, any and all subranges have a minimum value of 0 or more and a maximum value of 10 or less (e.g., 1 to 5). In some cases, the numerical values ​​presented as parameters may be negative values. In this case, negative values ​​such as -1, -2, -3, -10, -20, -30, etc. may be assumed as exemplary values ​​for the range set forth as "less than 10".

[0071] Although the present teachings have been illustrated with respect to one or more implementations, changes and / or modifications can be made to the illustrated embodiments without departing from the spirit and scope of the appended claims. For example, while a process is described as a sequence of acts or events, it will be understood that the present teachings are not limited by the ordering of such acts or events. Some acts may occur in different orders and / or may occur simultaneously with other acts or events than those described herein. Furthermore, not all process steps may be required to implement a methodology in accordance with one or more aspects or implementations of the present teachings. It will be understood that structural components and / or process steps may be added, or existing structural components and / or process steps may be removed or modified. Furthermore, one or more of the acts described herein may be performed in one or more other acts and / or phases. Still further, to the extent that "including," "includes," "having," "has," "with," or variations thereof are used in any of the detailed description and claims, these terms are intended to be as inclusive as the term "comprising." The phrase "at least one of" is used to mean that one or more of the listed items may be selected. The phrase "one or more of," as used herein in reference to a list of items (e.g., A and B), means A only, B only, or A and B. Additionally, in the description and claims herein, the term "on" as used in reference to two materials, i.e., one "on" the other, means that there is at least some contact between the materials, while "over" means that the materials are in close proximity, possibly with one or more additional intervening materials, such that contact is possible but not required.Neither "on" nor "over" as used herein implies any directionality. The term "conformal" describes a coating material where the angle of the underlying material is protected by the conformal material. The term "about" indicates that some variation may be made to the recited value without causing a process or structure incompatibility with the illustrated implementation. Finally, "exemplary" indicates that the description is used as an example, rather than implying ideality. Other implementations of the present teachings will be apparent to those skilled in the art from consideration of the specification and practice of the present disclosure. It is intended that the specification and examples be considered as exemplary only, with a true scope and spirit of the present teachings being indicated by the following claims.

[0072] Relative position terms used in this application are defined with respect to a plane parallel to the conventional plane or working surface of the workpiece, regardless of the orientation of the workpiece. The terms "horizontal" or "lateral" as used in this application are defined as a plane parallel to the conventional plane or working surface of the workpiece, regardless of the orientation of the workpiece. The term "vertical" refers to a direction perpendicular to the horizontal. Terms such as "on," "side (e.g., the side of a "sidewall")," "high," "low," "over," "top," and "below" are defined with respect to a conventional plane or working surface that is located on top of the workpiece, regardless of the orientation of the workpiece.

Claims

1. A composite panel (100), comprising: a composite panel segment (104-110) comprising a plurality of carbon fiber plies (200-206) disposed along a length and a width of the composite panel segment (104-110), the length of the composite panel segment (104-110) comprising a longitudinal axis (A); the plurality of carbon fiber plies (200-206) are stacked into a carbon fiber laminate that provides at least a portion of the thickness of the composite panel segment (104-110); the carbon fiber laminate comprises at least eight carbon fiber plies (200-206); each carbon fiber ply of the carbon fiber laminate has a stiffness, the stiffness of each carbon fiber ply being determined by the orientation of the carbon fibers in each carbon fiber ply relative to the longitudinal axis (A); the plurality of carbon fiber plies (200-206) of the carbon fiber laminate have a plurality of different orientations relative to the longitudinal axis (A), whereby the plurality of carbon fiber plies (200-206) of the carbon fiber laminate have a plurality of different stiffnesses; A composite panel (100) wherein the various orientations of the plurality of carbon fiber plies (200-206) of the carbon fiber laminate are self-symmetrical or generally self-symmetrical about a midpoint (400, 500, 600) of the carbon fiber laminate across a thickness of the composite panel segment (104-110), the midpoint (400, 500, 600) being parallel to the longitudinal axis (A), and the various orientations vary from being self-symmetrical for no more than two carbon fiber plies (200-206) when generally self-symmetrical.

2. the longitudinal axis (A) of the composite panel segments (104-110) is oriented at 0 degrees; 2. The composite panel (100) of claim 1, wherein the various orientations of the plurality of carbon fiber plies (200-206) include 0 degrees, 45 degrees, 90 degrees, and -45 degrees relative to the longitudinal axis (A) of the composite panel segment (104-110).

3. The composite panel (100) of claim 2, wherein the various orientations of the plurality of carbon fiber plies (200-206) of the carbon fiber laminate are self-symmetrical across a thickness of the composite panel segment (104-110).

4. the composite panel segments (104-110) include an intermediate segment (108) positioned between a base segment (106) and an upper segment (110); the intermediate segment (108) has a Poisson's ratio of 0.38 to 0.50; The composite panel (100) of claim 2 or 3, wherein between 46% and 58% of the plurality of carbon fiber plies (200-206) are oriented at 0 degrees.

5. 5. The composite panel (100) of any one of claims 1 to 4, wherein the plurality of carbon fiber plies (200-206) of the carbon fiber laminate comprises a plurality of repeatable groupings of carbon fiber plies (200-206), each repeatable grouping of carbon fiber plies (200-206) being repeated two, four, or six times.

6. the composite panel segments (104-110) include a mid-segment (108), the carbon fiber laminate is a mid-segment carbon fiber laminate, and the composite panel (100) comprises: a base segment (106) and an upper segment (110), the intermediate segment (108) being positioned between the base segment (106) and the upper segment (110); The base segment (106) a plurality of carbon fiber plies (202) disposed along a length and width of the base segment (106), the length of the base segment (106) being parallel to the longitudinal axis (A); the plurality of carbon fiber plies (202) of the base segment (106) are stacked into a base segment carbon fiber laminate that provides at least a portion of the thickness of the base segment (106); the base segment carbon fiber laminate comprising at least 10 carbon fiber plies (202); the plurality of carbon fiber plies (202) of the base segment carbon fiber laminate have a plurality of different orientations relative to the longitudinal axis (A), whereby the plurality of carbon fiber plies (202) of the base segment carbon fiber laminate have a plurality of different stiffnesses; 4. The composite panel (100) of claim 1, wherein the various orientations of the plurality of carbon fiber plies (200-206) of the base segment carbon fiber laminate are self-symmetrical or generally self-symmetrical across a thickness of the base segment (106) about a midpoint (600) of the base segment carbon fiber laminate, the midpoint (600) of the base segment being parallel to the longitudinal axis (A), and the various orientations of the plurality of carbon fiber plies (200-206) of the base segment carbon fiber laminate vary from being self-symmetrical for no more than two carbon fiber plies (200-206) when generally self-symmetrical.

7. the longitudinal axis (A) of the composite panel segments (104-110) is oriented at 0 degrees; the various orientations of the plurality of carbon fiber plies (204) of the intermediate segment carbon fiber laminate and the plurality of carbon fiber plies (202) of the base segment carbon fiber laminate include 0 degrees, 45 degrees, 90 degrees, and -45 degrees relative to the longitudinal axis (A) of the composite panel segment (104-110); the intermediate segment (108) has a Poisson's ratio of 0.38 to 0.50; The composite panel (100) of claim 6, wherein the base segment (106) and the top segment (110) each have a Poisson's ratio of between 0.36 and 0.

50.

8. The upper segment (110) a plurality of carbon fiber plies (206) disposed along a length and width of the upper segment (110), the length of the upper segment (110) being parallel to the longitudinal axis (A); the plurality of carbon fiber plies (206) of the upper segment (110) are stacked into an upper segment carbon fiber stack that provides at least a portion of the thickness of the upper segment (110); the upper segment carbon fiber laminate comprising at least 10 carbon fiber plies (206); the plurality of carbon fiber plies (206) of the upper segment carbon fiber laminate have a plurality of different orientations relative to the longitudinal axis (A), whereby the plurality of carbon fiber plies (206) of the upper segment carbon fiber laminate have a plurality of different stiffnesses; the various orientations of the plurality of carbon fiber plies (206) of the upper segment carbon fiber laminate are self-symmetrical or generally self-symmetrical around a midpoint (700) of the upper segment carbon fiber laminate across a thickness of the upper segment (110), the midpoint (700) of the upper segment carbon fiber laminate being parallel to the longitudinal axis (A), and the various orientations of the plurality of carbon fiber plies (206) of the upper segment carbon fiber laminate vary from being self-symmetrical in any two or less carbon fiber plies (206) when generally self-symmetrical; 8. The composite panel (100) of claim 6 or 7, wherein the plurality of carbon fiber plies (206) of the upper segment carbon fiber laminate are mirror images of the plurality of carbon fiber plies (202) of the base segment carbon fiber laminate.

9. the top segment carbon fiber laminate, the mid-segment carbon fiber laminate, and the base segment carbon fiber laminate are generally self-symmetrical about the midpoint (500) of the mid-segment carbon fiber laminate; 9. The composite panel (100) of claim 8, wherein the top segment carbon fiber laminate, the mid segment carbon fiber laminate, and the base segment carbon fiber laminate form at least a portion of a stringer (102) of the composite panel (100).

10. a skin segment (104), wherein the upper segment (110), the intermediate segment (108), and the base segment (106) are positioned on sides of the skin segment (104), and the base segment (106) is positioned between the intermediate segment (108) and the skin segment (104); The skin segment (104) a plurality of carbon fiber plies (200) disposed along a length and width of the skin segment (104), the length of the skin segment (104) being parallel to the longitudinal axis (A); the plurality of carbon fiber plies (200) of the skin segment (104) are stacked into a skin segment carbon fiber laminate that provides at least a portion of the thickness of the skin segment (104); the skin segment carbon fiber laminate comprising at least eight carbon fiber plies (200); the plurality of carbon fiber plies (200) of the skin segment carbon fiber laminate have a plurality of different orientations relative to the longitudinal axis (A), whereby the plurality of carbon fiber plies (200) of the skin segment carbon fiber laminate have a plurality of different stiffnesses; 10. The composite panel (100) of claim 8 or 9, wherein the various orientations of the plurality of carbon fiber plies (200) of the skin segment carbon fiber laminate are self-symmetrical about a midpoint (400) of the skin segment carbon fiber laminate through a thickness of the skin segment (104), the midpoint (400) of the skin segment carbon fiber laminate being parallel to the longitudinal axis (A).

11. the longitudinal axis (A) of the composite panel segments (104-110) is oriented at 0 degrees; the various orientations of the plurality of carbon fiber plies (204) of the intermediate segment carbon fiber laminate, the plurality of carbon fiber plies (202) of the base segment carbon fiber laminate, and the plurality of carbon fiber plies (200) of the skin segment carbon fiber laminate include 0 degrees, 45 degrees, 90 degrees, and -45 degrees relative to the longitudinal axis (A) of the composite panel segments (104-110); Between 46% and 58% of the plurality of carbon fiber plies (204) of the mid-segment carbon fiber laminate are oriented at 0 degrees; 11. The composite panel (100) of claim 10, wherein between 40% and 46% of the plurality of carbon fiber plies (206, 202) of each of the top segment carbon fiber laminate and the base segment carbon fiber laminate are oriented at 0 degrees.

12. 12. The composite panel (100) of claim 10 or 11, wherein the base segment (106) and the top segment (110) each have a Poisson's ratio of from 0.36 to 0.

50.

13. the top segment (110), the middle segment (108), and the base segment (106) form at least a portion of a solid laminate; The composite panel (100) of any one of claims 10 to 12, wherein a Poisson's ratio mismatch between the solid laminate and the skin segment (104) is in the range of -0.06 to 0.

06.

14. A method for forming a composite panel (100), comprising: laminating together a plurality of carbon fiber plies (200-206) to form a carbon fiber laminate of a composite panel segment (104-110), said carbon fiber laminate comprising at least ten carbon fiber plies (200-206); bonding a plurality of said carbon fiber plies (200-206) together Including, the plurality of carbon fiber plies (200-206) are disposed along the length and width of the composite panel segment (104-110); said plurality of carbon fiber plies (200-206) forming at least a portion of the thickness of said composite panel segment (104-110); the length of the composite panel segment (104-110) includes a longitudinal axis (A); Each carbon fiber ply of the carbon fiber laminate has a stiffness; the stiffness of each carbon fiber ply is determined by the orientation of the carbon fibers in each carbon fiber ply relative to the longitudinal axis (A); the plurality of carbon fiber plies (200-206) of the carbon fiber laminate have a plurality of different orientations relative to the longitudinal axis (A), whereby the plurality of carbon fiber plies (200-206) of the carbon fiber laminate have a plurality of different stiffnesses; the various orientations of the plurality of carbon fiber plies (200-206) of the carbon fiber laminate are self-symmetrical or generally self-symmetrical about a midpoint (400, 500, 600, 700) of the carbon fiber laminate through a thickness of the composite panel segment (104-110); said intermediate points (400, 500, 600, 700) are parallel to said longitudinal axis (A); The method wherein said various orientations vary from being self-symmetrical in no more than two carbon fiber plies (200-206) when being generally self-symmetrical.

15. orienting the longitudinal axis (A) of the composite panel segments (104-110) at 0 degrees; orienting each carbon fiber ply of the plurality of carbon fiber plies (200-206) to have an orientation at one of 0 degrees, 45 degrees, 90 degrees, and -45 degrees relative to the longitudinal axis (A); stacking at least ten carbon fiber plies (200-206) such that the various orientations of the plurality of carbon fiber plies (200-206) of the carbon fiber laminate are self-symmetrical; The method of claim 14 further comprising:

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