Laminate Stack for Solid Laminated Reinforcement Plates

Abstract

The name of the present invention is a ply stack for a solid laminated reinforcement panel. The composite panel includes a plurality of segments, where each segment includes a plurality of reinforcing plies, such as carbon fiber plies. The collimated fiber bands (e.g., carbon fibers) within each reinforcing ply are oriented in a single direction, which can be, for example, 0°, 45°, 90°, and -45°. Each segment can include a stack of reinforcing plies, where the collimated fiber bands of each reinforcing ply are oriented in one of these directions. The orientation of the collimated fiber bands forming the reinforcing plies in the stack determines the stiffness of the stack and the stiffness of the segments including the stack. The stiffness of each segment is controlled to reduce the stiffness mismatch between adjacent segments forming the composite panel, thereby reducing the separation of the segments during use.

Description

Technical Field

[0001] In general, the present teachings relate to structural components such as composite panels for manufacturing aircraft or other articles, and more particularly to composite structural component assemblies and methods of forming the same. Background Art

[0002] The fuselage, wings, tail, empennage, etc. of an aircraft can be manufactured using composite panels such as solid laminated spars attached to a support frame. The composite panel can include an outer skin and an internal laminate, where the internal laminate at least partially forms the spar. The skin and the internal laminate can each be formed of a plurality of carbon fiber layers or plies bonded with a thermoplastic resin adhesive or matrix (referred to as prepreg). The skin of the composite panel is exposed and provides the outer surface of the aircraft during use. The composite panel must be able to resist damage from stresses and strains caused, for example, by aerodynamic loads and temperature changes placed on the composite panel during use. These stresses and strains can cause the composite panel to buckle, bend, twist, expand, contract, etc. during use, and over time, can cause delamination or other damage to the carbon fiber plies.

[0003] A composite panel that is more resistant to damage caused by stresses and strains placed on the composite panel during use would be welcome in the art. Summary of the Invention

[0004] The following presents a simplified summary of the invention in order to provide a basic understanding of some aspects of one or more embodiments of the present teachings. This summary is not an extensive overview, nor is it intended to identify key or critical elements of the present teachings, nor is it intended 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 that is presented later.

[0005] In an implementation, a composite panel includes a composite panel segment that can include a plurality of carbon fiber plies arranged along the length and width of the segment, where the length of the composite panel segment has a longitudinal axis. The plurality of carbon fiber plies are laminated into a carbon fiber stack that provides at least a portion of the thickness of the composite panel segment and includes at least 8 carbon fiber plies. Each carbon fiber ply of the carbon fiber stack has a stiffness, where the stiffness of each carbon fiber ply is determined by the orientation of the carbon fibers within 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, and thus, the plurality of carbon fiber plies of the carbon fiber stack can include a plurality of different stiffnesses. The different orientations of the plurality of carbon fiber plies of the carbon fiber stack are self-symmetric or mostly self-symmetric about the midpoint of the carbon fiber stack through the thickness of the composite panel segment, where the midpoint is parallel to the longitudinal axis, where, when mostly self-symmetric, the different orientations are self-symmetrically different in no more than two carbon fiber plies.

[0006] Optionally, the longitudinal axis of the composite panel segment can be oriented at 0°, and the different orientations of the plurality of carbon fiber plies include 0°, 45°, 90°, and -45° relative to the longitudinal axis of the composite panel segment. The different orientations of the plurality of carbon fiber plies in the carbon fiber stack can be self-symmetric through the thickness of the composite panel segment. Additionally, the composite panel segment can be an intermediate segment located between a base segment and a top segment, and 46% to 58% of the plurality of carbon fiber plies can be oriented at 0°. The intermediate segment can have an axial elastic modulus of from about 11.4 megapounds per square inch (msi) to about 13.4 msi. Additionally, the intermediate segment can have a Poisson's ratio of from about 0.38 to 0.50.

[0007] In an implementation, the plurality of carbon fiber plies in the carbon fiber stack can include a plurality of repeatable groups of carbon fiber plies, where each repeatable group of carbon fiber plies is repeated 2, 4, or 6 times. The different orientations of the plurality of carbon fiber plies in the carbon fiber stack can be substantially self-symmetric through the thickness of the composite panel segment. The carbon fiber stack can include at least 16 carbon fiber plies.

[0008] In an implementation, the composite panel segment can be an intermediate segment, the carbon fiber stack can be an intermediate segment carbon fiber stack, and the composite panel can further include a base segment and a top segment, where the intermediate segment is located between the base segment and the top segment. In this implementation, the base segment can include a plurality of carbon fiber plies disposed along the length and width of the base segment, and the length of the base segment is parallel to the longitudinal axis. The plurality of carbon fiber plies of the base segment are laminated into a base segment carbon fiber stack that provides at least a portion of the thickness of the base segment, and the base segment carbon fiber stack can include at least 10 carbon fiber plies. Additionally, the plurality of carbon fiber plies of the base segment carbon fiber stack can have a plurality of different orientations relative to the longitudinal axis, and thus, the plurality of carbon fiber plies of the base segment carbon fiber stack can include a plurality of different stiffnesses, and the different orientations of the plurality of carbon fiber plies of the base segment carbon fiber stack are self-symmetric or substantially self-symmetric about the midpoint of the base segment carbon fiber stack through the thickness of the base segment, where the midpoint of the base segment is parallel to the longitudinal axis, where, when substantially self-symmetric, the different orientations of the plurality of carbon fiber plies of the base segment carbon fiber stack are self-symmetrically different in no more than two carbon fiber plies. Optionally, the longitudinal axis of the composite panel segment is oriented at 0°, and the different orientations of the plurality of carbon fiber plies of the intermediate segment carbon fiber stack and the plurality of carbon fiber plies of the base segment carbon fiber stack include 0°, 45°, 90°, and -45° relative to the longitudinal axis of the composite panel segment. The intermediate segment can have an axial elastic modulus of from about 11.4 megapounds per square inch (msi) to about 13.4 msi, while the base segment and the top segment can each have an axial elastic modulus of from about 10.0 msi to about 11.4 msi.

[0009] In an implementation, the middle section may have a Poisson's ratio of about 0.38 to about 0.50, and the base section and the top section may each have a Poisson's ratio of about 0.36 to about 0.50. The top section may include a plurality of carbon fiber plies arranged along the length and width of the top section, and the length of the top section may be parallel to the longitudinal axis. The plurality of carbon fiber plies of the top section may be laminated into a top section carbon fiber stack that provides at least a portion of the thickness of the top section. Additionally, the top section carbon fiber stack may include at least 10 carbon fiber plies, the plurality of carbon fiber plies of the top section carbon fiber stack may include a plurality of different orientations relative to the longitudinal axis, and thus, the plurality of carbon fiber plies of the top section carbon fiber stack may include a plurality of different stiffnesses. The different orientations of the plurality of carbon fiber plies of the top section carbon fiber stack may be self-symmetric or substantially self-symmetric about the midpoint of the top section carbon fiber stack through the thickness of the top section, wherein the midpoint of the top section carbon fiber stack may be parallel to the longitudinal axis. When substantially self-symmetric, the different orientations of the plurality of carbon fiber plies of the top section carbon fiber stack may be self-symmetrically different in no more than two carbon fiber plies, and the plurality of carbon fiber plies of the top section carbon fiber stack may be mirror-symmetric with the plurality of carbon fiber plies of the base section carbon fiber stack.

[0010] In an optional implementation, the top section carbon fiber stack, the middle section carbon fiber stack, and the base section carbon fiber stack may be substantially self-symmetric about the midpoint of the middle section carbon fiber stack. Additionally, the top section carbon fiber stack, the middle section carbon fiber stack, and the base section carbon fiber stack may form at least a portion of a spar of a composite panel. The composite panel may further include a skin section, wherein the top section, the middle section, and the base section may be located on one side of the skin section, and the base section may be located between the middle section and the skin section. Additionally, the skin section may include a plurality of carbon fiber plies arranged along the length and width of the skin section, and the length of the skin section may be parallel to the longitudinal axis. The plurality of carbon fiber plies of the skin section may be laminated into a skin section carbon fiber stack that provides at least a portion of the thickness of the skin section. Additionally, the skin section carbon fiber stack may include at least 8 carbon fiber plies, the plurality of carbon fiber plies of the skin section carbon fiber stack may include a plurality of different orientations relative to the longitudinal axis, and thus, the plurality of carbon fiber plies of the skin section carbon fiber stack may include a plurality of different stiffnesses. The different orientations of the plurality of carbon fiber plies of the skin section carbon fiber stack may be self-symmetric about the midpoint of the skin section carbon fiber stack through the thickness of the skin section, wherein the midpoint of the skin section carbon fiber stack is parallel to the longitudinal axis.

[0011] In practice, the longitudinal axis of the composite panel segment is oriented at 0°, and the different orientations of the plurality of carbon fiber laminae of the carbon fiber stack in the intermediate segment, the plurality of carbon fiber laminae of the carbon fiber stack in the base segment, and the plurality of carbon fiber laminae of the carbon fiber stack in the skin segment may include 0°, 45°, 90°, and -45° relative to the longitudinal axis of the composite panel segment, and approximately 46% to approximately 58% of the plurality of carbon fiber laminae of the carbon fiber stack in the intermediate segment may be oriented at 0°. Optionally, approximately 40% to approximately 46% of the plurality of carbon fiber laminae of each of the carbon fiber stacks in the top segment carbon fiber stack and the base segment carbon fiber stack are oriented at 0°. Further optionally, the intermediate segment may have an axial elastic modulus of approximately 11.4 million pounds per square inch (msi) to approximately 13.4 msi, the base segment and the top segment may each have an axial elastic modulus of approximately 10.0 msi to approximately 11.4 msi, and the skin segment may have an axial elastic modulus of approximately 7.7 msi to approximately 10.0 msi.

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

[0013] In another implementation, the composite panel includes a composite panel segment having a plurality of reinforcing laminae disposed along the length and width of the segment, wherein the length of the composite panel segment includes a longitudinal axis. The plurality of reinforcing laminae are laminated into a reinforcing lamina stack that provides at least a part of the thickness of the composite panel segment. The plurality of reinforcing laminae of the reinforcing lamina stack include a plurality of different orientations relative to the longitudinal axis, and the different orientations of the plurality of reinforcing laminae of the reinforcing lamina stack are self-symmetric or substantially self-symmetric about the midpoint of the reinforcing lamina stack through the thickness of the composite panel segment, wherein the midpoint is parallel to the longitudinal axis, and wherein when substantially self-symmetric, the different orientations are symmetrically different in no more than two carbon fiber laminae. Optionally, the plurality of reinforcing laminae includes at least 16 reinforcing laminae, and the different orientations of the plurality of reinforcing laminae are self-symmetric.

[0014] In another embodiment, a method for forming a composite panel includes: stacking a plurality of carbon fiber plies together to form a carbon fiber stack of a composite panel segment, wherein the carbon fiber stack includes at least 10 carbon fiber plies; and bonding the plurality of carbon fiber plies together. In this embodiment, the plurality of carbon fiber plies are arranged along the length and width of the composite panel segment, the plurality of carbon fiber plies form at least a part of the thickness of the composite panel segment, and the length of the composite panel segment includes a longitudinal axis. Further, each carbon fiber ply of the carbon plate stack has a stiffness, the stiffness of each carbon fiber ply is determined by the orientation of the carbon fibers within each carbon fiber ply relative to the longitudinal axis, and the plurality of carbon fiber plies of the carbon fiber stack have a plurality of different orientations relative to the longitudinal axis, and thus, the plurality of carbon fiber plies of the carbon fiber stack include a plurality of different stiffnesses. Additionally, the different orientations of the plurality of carbon fiber plies of the carbon fiber stack are self-symmetric or substantially self-symmetric about the midpoint of the carbon fiber stack through the thickness of the composite panel segment, and the midpoint is parallel to the longitudinal axis. When substantially self-symmetric, the different orientations are symmetrically different in no more than two carbon fiber plies. The method may optionally include orienting the longitudinal axis of the composite panel segment to 0°, orienting each of the plurality of carbon fiber plies to have an orientation of one of 0°, 45°, 90°, and -45° relative to the longitudinal axis, and stacking at least 10 carbon fiber plies such that the different orientations of the plurality of carbon fiber plies of the carbon fiber stack are self-symmetric. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] The drawings incorporated in and constituting a part of this specification illustrate embodiments of the present teachings and, together with the specification, are used to explain the principles of the disclosure. In the figures:

[0016] Figure 1 is a perspective view of a composite panel including a skin and stringers, implemented according to an example of the present teachings.

[0017] Figure 2 is a cross-section taken along Figure 1 2-2.

[0018] Figure 3 is a plan view of four collimated fiber bands of four different reinforcing plies, implemented according to an embodiment of the present teachings.

[0019] Figure 4 is a table listing the reinforcing plies and their fiber orientations in a reinforcing ply stack of a skin segment of a composite panel, implemented according to an example of the present teachings.

[0020] Figure 5 is a table listing the reinforcing plies and their fiber orientations in a reinforcing ply stack of an intermediate segment that partially provides the stringers of the composite panel, implemented according to an example of the present teachings.

[0021] Figure 6A table listing the reinforcing plies and their fiber orientations in the reinforcing ply stack of the base section implemented according to an example of the present teachings, the base section partially providing the spar of the composite panel.

[0022] Figure 7 A table listing the reinforcing plies and their fiber orientations in the reinforcing ply stack of the top section implemented according to an example of the present teachings, the top section partially providing the spar of the composite panel.

[0023] Figure 8 Depicts the runout of the spar including the base section and the top section in an example implementation of the present teachings.

[0024] Figure 9 A table listing the parameters and characteristics of the individual sub-components of the composite panel implemented according to the present teachings.

[0025] Figure 10 A graph depicting the axial elastic modulus (i.e., stiffness) of the individual components of the composite panel in an example implementation of the present teachings.

[0026] Figure 11 A schematic diagram of the formation of the spar of the composite panel in an example implementation of the present teachings.

[0027] It should be noted that some details of the drawings have been simplified and drawn to facilitate the understanding of the present teachings, rather than maintaining strict structural accuracy, details, and proportions. Detailed Description

[0028] Reference will now be made in detail to the exemplary implementations of the present teachings, examples of which are shown in the drawings. Generally and / or when convenient, the same reference numerals will be used throughout the drawings to refer to the same or similar components.

[0029] Implementations of the present teachings include a composite panel and a method for manufacturing the composite panel. The composite panel can be a sub-component of a larger structure such as an aircraft, aerospace vehicle, or another structure. The composite panel can include a plurality of sections, each section including a plurality of stacked reinforcing material plies (i.e., reinforcing plies), which can be or include, for example, carbon fibers. In an implementation, the configuration of the composite panel as designed has increased delamination resistance and thermal cracking resistance compared to some conventional composite panels, which delamination and thermal cracking can be caused by thermal mismatch, bending or warping, or other stresses and strains during use of the composite panel. The composite panel can include solid laminated spars and a configuration that contributes to the spar runout.

[0030] Figure 1 Is a perspective view of a portion of a composite panel 100 implemented according to the present teachings, and Figure 2 Is along Figure 1The cross-section taken along line 2-2. For purposes of description, the composite panel 100 includes solid laminated stringers (hereinafter referred to as "stringers") 102 and skin segments (i.e., skin laminates) 104, but other types of composite panels that do not include stringers 102 and / or skin segments 104 may be contemplated. Stringers and skins are well known in, for example, aircraft manufacturing and other industrial fields.

[0031] Figure 1 The stringer 102 includes a base segment 106, an intermediate segment 108, and a top segment 110. As Figure 1 and Figure 2 depicted, the base segment 106 and the intermediate segment 108 are located between the skin segment 104 and the top segment 110, and the intermediate segment 108 is located between the base segment 106 and the top segment 110.

[0032] As described in more detail below, each of the skin segment 104, the base segment 106, the intermediate segment 108, and the top segment 110 respectively includes a plurality of reinforcing plies 200, 202, 204, and 206. Although, for simplicity, only some of the reinforcing plies 200-206 that only partially extend across each segment 104-110 are depicted, it should be understood that each segment 104-110 includes at least 8 reinforcing plies 200-206 that extend across most or all of each segment 104-110. In some configurations, 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 may be, for example, a plurality of carbon fiber plies. The plurality of reinforcing plies may be pre-impregnated with an adhesive resin such as a thermoplastic resin, which bonds the plurality of reinforcing plies of each segment together and may be referred to as 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 depicted. Each of the plurality of plies 200-208 is laminated (i.e., stacked) along the thickness of each segment 104-110 to form a reinforcing ply stack (e.g., a carbon fiber stack). In Figure 2 the orientation of the illustration, each reinforcing ply extends in the horizontal direction, wherein each reinforcing ply stack includes a plurality of vertically stacked reinforcing plies.

[0033] Each of segments 104 - 110 can be formed using an automated fiber placement (AFP) process, where a computer-controlled fiber placement head of a fiber placement system deposits a plurality of prepreg tows that form a collimated fiber tape onto a surface. In some processes, the collimated fiber tape is deposited onto a mandrel that provides or defines the shape of the final component to be produced. In these processes, segments 104 - 110 are chemically bonded and / or mechanically fastened together. In other processes, the skin segment 104 can be deposited and fabricated first, and then the base segment 106 can be deposited onto the skin segment 104, the intermediate segment 108 can be deposited onto the base segment 106, and the top segment 110 can be deposited onto the intermediate segment 108. In an alternative, the order can be reversed such that the top segment 110 is deposited first and the skin segment 104 is deposited onto the base segment 106 last.

[0034] Figure 1 A longitudinal axis “A” is depicted that extends through and is parallel to the length of the composite panel 100 and the length of the stringer 102. For the purposes of the present disclosure, the longitudinal axis A that extends through and is parallel to the length of the composite panel 100 is defined as 0° relative to the composite panel 100 (i.e., in the axial direction). During the deposition of each collimated fiber tape, the fiber placement head can travel in the 0° direction to deposit the collimated fiber tape in the 0° direction such that the collimated fiber tape is oriented at 0°. In this orientation, the fiber placement head requires the fewest number of passes across the composite panel to deposit the plurality of collimated fiber tapes that form a single reinforcing lamina.

[0035] Figure 3 is a schematic plan view depicting the composite panel 100, where the length “L” of the composite panel 100 and thus the axis “A” are oriented from top to bottom at 0°, and the width “W” is oriented from left to right. During the deposition of each collimated fiber tape 300 - 306, the travel direction of the fiber placement head can be changed to deposit the collimated fiber tapes 300 - 306 in various different directions. Figure 3 Schematically depicts four collimated fiber tapes 300 - 306 deposited during four separate passes of the fiber placement head in four different directions to form four different laminae. The collimated fiber tapes 300, 302, 304, and 306 are oriented at 0°, 45°, 90°, and - 45°, respectively. Although for simplicity, Figure 3 depicts the result of only one pass of the fiber placement head for each reinforcing lamina, it will be understood that a sufficient number of passes of the fiber placement head are made to complete each reinforcing lamina. In addition, the four collimated fiber tapes 300 - 306 and thus their respective reinforcing laminae are stacked vertically (i.e., perpendicular to the plane of the page).

[0036] The specific orientation of the collimated fiber bands 300-306 that stack to form one of the segments 104-110 will determine the stiffness of the segments 104-110. For example, a segment including multiple reinforcing plies with collimated fiber bands oriented only at 0° will be a relatively stiff segment, while a segment including multiple reinforcing plies oriented only at 90° will be a relatively flexible segment.

[0037] This teaching includes a composite panel having multiple segments. During the design and manufacture of the composite panel, the stiffness of each segment is controlled by the selection of the orientation of the collimated fiber bands within each of the multiple reinforcing plies that make up the segment. Additionally, compared to some conventional designs, the orientation of the collimated fiber bands of the reinforcing plies within each segment is designed and selected to improve the adhesion between adjacent reinforcing plies and thus improve the overall adhesion of each segment to reduce delamination within each segment. Additionally, the stiffness of each of the multiple segments making up the composite panel is designed to improve the adhesion between adjacent segments and thus reduce the stiffness mismatch between adjacent segments to reduce separation of the composite panel between adjacent segments.

[0038] In an implementation of this teaching, the orientation of the multiple reinforcing plies within each individual segment across the thickness of each segment is designed, controlled, and / or selected to be self-symmetric or substantially self-symmetric. For the purposes of this disclosure, unless otherwise stated, in a self-symmetric segment, the entirety of the multiple reinforcing plies that make up the self-symmetric segment is mirror symmetric about the midpoint of the thickness of the self-symmetric segment. In a substantially self-symmetric segment, no more than two of the multiple reinforcing plies that make up the substantially self-symmetric segment are out of order or different from mirror symmetry about the midpoint of the thickness of the segment. In other words, in a substantially self-symmetric segment, changing the orientation of no more than two reinforcing plies (i.e., one ply or two plies) will result in the segment being self-symmetric.

[0039] Furthermore, as used herein, when referring to the multiple reinforcing plies that form the entire thickness of two or more segments, the term "self-symmetric" indicates that the multiple reinforcing plies that form two or more segments are mirror symmetric about the midpoint of the two or more segments, while "substantially self-symmetric" indicates that among the multiple reinforcing plies that form two or more segments, no more than two of the multiple reinforcing plies that make up the substantially self-symmetric stack are out of order or different from mirror symmetry about the midpoint of the thickness of the two or more segments. In other words, across two or more segments, changing the orientation of no more than two reinforcing plies (i.e., one ply or two plies) will result in the multiple reinforcing plies that form two or more segments being self-symmetric. It will be understood that two or more self-symmetric or substantially self-symmetric segments may be placed adjacent to each other or may have other structures, such as other segments located between them.

[0040] In addition, the "midpoint" of a segment (or two or more segments) is the midpoint relative to the number of reinforcing laminae that make up the reinforcing lamina stack. In a segment (or two or more segments) that includes an odd number of reinforcing laminae, the midpoint is the middle reinforcing lamina; in a segment (or two or more segments) that includes an even number of reinforcing laminae, the midpoint will be between two reinforcing laminae at the middle of the segment (or the middle of two or more segments), where an equal number of reinforcing laminae are on opposite sides of the midpoint. Similarly, "corresponding" reinforcing laminae refer to two reinforcing laminae that are equidistant from a common midpoint but on opposite sides of it.

[0041] As described above, in an implementation of the present teachings, all of the reinforcing laminae of two or more segments that make up a composite panel can be self-symmetric or substantially self-symmetric across the thickness of the composite panel that includes two or more segments. With respect to Figure 1 and Figure 2 an example, the three segments 106-110 that make up the spar 102 can be self-symmetric or substantially self-symmetric. In addition, the stiffness of the spar 102 (i.e., the segments 106-110 that make up the spar 102) can be matched to the stiffness of the skin segment 104. By matching the stiffness of these two structures (i.e., the skin segment 104 and the spar 102), stiffness mismatches that can cause the spar to separate from the skin can be reduced or eliminated.

[0042] Figures 4 - 7 is a table listing example reinforcing lamina orientations for the four segments 104-110 used for Figure 1 where Figure 4 depicts an example reinforcing lamina orientation for the skin segment 104, Figure 5 depicts an example reinforcing lamina orientation for the middle segment 108, Figure 6 depicts an example reinforcing lamina orientation for the base segment 106, and Figure 7 depicts an example reinforcing lamina orientation for the top segment 110. As Figures 4 - 7 depicted, the skin segment 104 includes 40 reinforcing laminae, the middle segment 108 includes 32 reinforcing laminae, and each of the base segment 106 and the top segment 110 includes 14 laminae. It should be understood that each of the segments 104-110 can include more or fewer laminae than those depicted. For example, in an implementation, each of the segments 104-110 can include at least 8 laminae and can include more than 1000 laminae. In each of these implementations, the reinforcing laminae within each segment will be self-symmetric or substantially self-symmetric. In addition, the reinforcing laminae across all segments that make up the composite panel will be self-symmetric or substantially self-symmetric.

[0043] Figure 4 is a table listing in a composite panel having 40 reinforcing laminae 200 (in Figure 2Table of the reinforcing plies and their fiber orientations in the self-symmetric skin section 104 depicted in. As Figure 4 depicted in, the bottommost ply (ply #1) is oriented in the same direction (45°) as the topmost ply (ply #40), ply #2 is oriented in the same direction (90°) as ply #39, and so on, until the midpoint 400 between reinforcing plies #20 and #21, where both reinforcing plies #20 and #21 are oriented at 0°. Additionally, the skin section 104 may include a reinforcing ply 200 that consists of approximately 25% to approximately 40% collimated fiber bands oriented in the 0° direction. Using these parameters, the skin section 104 will have an axial elastic modulus of approximately 7.7 megapounds per square inch (msi) to approximately 10.0 msi and a Poisson's ratio of approximately 0.32 to approximately 0.48.

[0044] Figure 5 is a table listing the reinforcing plies and their fiber orientations in the self-symmetric intermediate section 108 having 32 reinforcing plies 204 (depicted in Figure 2 ). As Figure 5 depicted in, the bottommost ply (ply #1) is oriented in the same direction (0°) as the topmost ply (ply #32), ply #2 is oriented in the same direction (0°) as ply #32, and so on, until the midpoint 500 between reinforcing plies #16 and #17, where both reinforcing plies #20 and #21 are oriented at 0°. Figure 5 The ply stack depicted in provides a repeatable group of reinforcing plies that can be replicated or repeated any even number of times (i.e., 2 times, 4 times, 6 times, etc.) to form a thicker intermediate section 108. This allows the entire solid laminate (e.g., spar 102) adjacent to the skin section 104 to have a thickness that can be selected or designed based on the expected loads, deformations, and damage protection during use. Additionally, the intermediate section 108 may include a reinforcing ply 204 that consists of approximately 46% to approximately 58% collimated fiber bands oriented in the 0° direction. Using these parameters, the intermediate section will have an axial elastic modulus of approximately 11.4 msi to approximately 13.4 msi and a Poisson's ratio of approximately 0.38 to approximately 0.50.

[0045] Figure 6 is a table listing the reinforcing plies and their fiber orientations in the substantially self-symmetric base section 106 having 14 reinforcing plies 202 (depicted in Figure 2 ). As Figure 6 depicted in, the bottommost ply (ply #1) is oriented in the same direction (45°) as the topmost ply (ply #14), ply #2 is oriented in the same direction (90°) as ply #13, and so on, until the asymmetric corresponding reinforcing plies 6 (oriented at 45°) and 9 (oriented at -45°). The remaining reinforcing plies 7 and 8 are symmetric about the midpoint 600 therebetween. Thus, the base section 106 (inFigure 2 Depicted in [the figure] is substantially self-symmetric because changing one lamina (i.e., changing lamina 6 to -45° or lamina 9 to 45°) will render the base segment 106 self-symmetric. Additionally, the base segment 106 may include a reinforcing lamina 202, which is composed of approximately 40% to approximately 46% of collimated fiber bands oriented in the 0° direction. Using these parameters, the base segment 106 will have an axial elastic modulus of approximately 10.0 msi to approximately 11.4 msi and a Poisson's ratio of approximately 0.36 to approximately 0.50.

[0046] Figure 7 is a table listing the reinforcing laminae and their fiber orientations in the substantially self-symmetric top segment 110 having 14 reinforcing laminae 206 (depicted in Figure 2 [the figure]). As depicted in Figure 7 [the figure], the bottommost lamina (lamina #N - 13) is oriented in the same direction (45°) as the topmost lamina (lamina #N), lamina #N - 1 is oriented in the same direction (90°) as lamina #N - 12, and so on, until the asymmetric corresponding reinforcing laminae N - 8 (oriented at -45°) and N - 5 (oriented at 45°). The remaining reinforcing laminae N - 7 and N - 6 are symmetric about the midpoint 700 therebetween. Thus, the top segment 110 ( Figure 2 ) is substantially self-symmetric because changing one lamina (i.e., changing lamina N - 8 to 45° or lamina N - 5 to -45°) will render the top segment 110 symmetric. Additionally, the top segment 110 may include a reinforcing lamina 206, which is composed of approximately 40% to approximately 46% of collimated fiber bands oriented in the 0° direction. Using these parameters, the top segment 110 will have an axial elastic modulus of approximately 10.0 msi to approximately 11.4 msi and a Poisson's ratio of approximately 0.36 to approximately 0.50.

[0047] Furthermore, segments 104 - 110 will include a reinforcing lamina composed of approximately 12% or more of collimated fiber bands oriented at a transverse angle (i.e., 90°).

[0048] Forming the skin segments 104, base segments 106, intermediate segments 108, and top segments 110 with the percentages of collimated fiber bands listed above will result in these segments having the axial elastic moduli and Poisson's ratios as described above. Forming segments with a specific range of axial elastic moduli and Poisson's ratios as described above will result in an integrated structure with high delamination resistance, reduced overall waviness, reduced weight, and reduced manufacturing and material costs compared to traditional structures formed using traditional methods. Forming segments 104 - 110 with an axial elastic modulus lower than that described above will result in a structure with increased weight, manufacturing time, and cost. Forming segments 104 - 110 with an axial elastic modulus higher than that described above will result in a structure with an increased tendency to delaminate due to increased mismatch stress between the skin segment 104 and the stringer 102. Forming segments 104 - 110 with a Poisson's ratio lower than that described above will result in a structure with reduced delamination resistance due to Poisson's ratio mismatch between the structures. Forming segments 104 - 110 with a Poisson's ratio higher than that described above will result in increased Poisson's ratio mismatch, reduced delamination resistance, and increased overall waviness.

[0049] Additionally, as Figure 6 and Figure 7 depicted, the base segment 106 and the top segment 110 are self - symmetric (i.e., mirror - symmetric) with respect to each other about the mid - point 650 therebetween. Thus, in the solid laminate end region or tip (such as the tip 800 depicted in Figure 8 ), for example, at the end of the stringer 102, the base segment 106 and the top segment 110 combine to form the self - symmetric tip 800. For example, during the formation of the stringer 102, the plurality of laminae 204 forming the intermediate segment 108 can be formed with different lengths such that the intermediate segment 108 becomes thinner and thinner until at the end of the stringer 102, the tip 800 includes only the base segment 106 and the top segment 110. In other words, the laminae constituting the self - symmetric reinforcing lamina stack of the intermediate segment 108 drop off in pairs of corresponding reinforcing laminae (e.g., lamina 1 and 32, then lamina 2 and 31, etc.) until only the base segment 106 and the top segment 110 remain to form the tip 800. As Figure 8 depicted, the base segment 106 and the top segment 110 have the same thickness. Additionally, as described above, the base segment 106 is mirror - symmetric with respect to the top segment 110 about the mid - point 802 at the interface between the two segments 106, 110. It should be understood that the mid - point 802 is similar to the mid - point 650 of Figure 6 .

[0050] Furthermore, because the intermediate segment 108 is self - symmetric about the mid - point 500, the stringer 102 formed by the base segment 106, the intermediate segment 108, and the top segment 110 is also self - symmetric, with the intermediate segment 108 located between the base segment 106 and the top segment 110.

[0051] Figure 9 shows the axial fiber percentage (i.e., the percentage of collimated fiber bands oriented at 0°), the transverse fiber percentage (i.e., the percentage of collimated fiber bands oriented at 90°), the axial elastic modulus, the Poisson's ratio, and the Poisson's ratio mismatch relative to the skin for each of the segments 104 - 110 in the example embodiments described herein. The Poisson's ratio mismatch between the skin segment 104 and the base segment 106 is in the range of approximately -0.06 to approximately 0.06. In addition, the Poisson's ratio mismatches between the base segment 106 and the intermediate segment, and between the top segment 110 and the intermediate segment 108 are also in the range of approximately -0.06 to approximately 0.06. These relatively low Poisson's ratio mismatches reduce or prevent delamination of the stringer 102 from the skin segment 104, and also reduce or prevent delamination of the base segment 106 and the top segment 110 from the intermediate segment 108.

[0052] In Figure 1 the composite panel 100, the overall stiffness of the composite panel gradually varies from the relatively soft skin segment 104 to the relatively stiff base segment 106 and top segment 110, to the much stiffer intermediate segment 108, where the base segment 106 and the top segment 110 are stiffer than the skin segment 104 but not as stiff as the intermediate segment 108. In an example embodiment, the stiffness of each of the segments 104 - 110 is depicted in Figure 10 which Figure 10 shows the gradual change in the axial elastic modulus between the segments. This gradual change reduces the stiffness mismatch between adjacent segments, which improves delamination resistance, thereby reducing or eliminating separation and delamination between the segments.

[0053] The overall combination and sequence of the proposed reinforcing ply stacks from the skin segment 104 to the segments 106 - 110 forming the stringer 102 can reduce the overall waviness of the reinforced composite panel 100 and can improve the build quality of the composite panel. In an implementation, compared to composite panels manufactured using traditional techniques, the overall waviness of the composite panel 100 can be reduced by 30% or more, or by 40% or more. This reduction in the overall waviness of the composite panel can result in less assembly work / rework. In addition, for example, by reducing work-related injuries caused by working in a narrow workspace during assembly, the reduction in overall waviness can reduce recurring costs and improve labor safety. As the overall waviness is reduced, the requirement to fill the reduced gaps between parts or structures due to the overall waviness is alleviated or eliminated.

[0054] As described above, Figure 3The collimated fiber bands 300 - 306 are deposited using a fiber placement head of a fiber placement system. The fiber placement head passes over the mandrel multiple times to deposit each reinforcement ply. When the first ply is complete, the fiber placement head can deposit the second ply over the first ply. The reinforcement ply with a 0° orientation requires the least deposition time because the 0° orientation requires the least number of passes of the fiber placement head over the mandrel to complete the reinforcement ply. In contrast, the reinforcement ply with a 90° orientation requires the most deposition time because the 90° orientation requires the most number of passes of the fiber placement head over the mandrel to complete the reinforcement ply.

[0055] In an implementation of this teaching, common multi - ply components or groups of reinforcement plies can be pre - assembled (i.e., pre - fabricated) and stored for later use. Then, various required multi - ply components can be removed from storage and assembled into the final component.

[0056] For example, Figure 11 Depicted are multiple multi - ply groups 1100, where each multi - ply group is labeled with letters A - J (i.e., 10 different multi - ply groups). Figure 11 Further depicted is a spar 102 that includes a base section 106, an intermediate section 108, and a top section 110. Multiple of each multi - ply group A - J can be pre - assembled and stored until the composite panel 100 including the spar 102 is fabricated. Before constructing the composite panel 100, the required number of each multi - ply group A - J is requested and transferred to the construction location. Construction begins by placing the first required multi - ply group of the spar 102 on the mandrel 1102, which in this example implementation is the three - ply group A of the base section 106. The next required reinforcement ply orientation is 0°, which, as described above, can be deposited relatively quickly. Thus, the reinforcement ply with a 0° orientation is deposited using the fiber placement head 1104 of the fiber placement system. Next, each required multi - ply group A - J is placed, and each required reinforcement ply with a 0° orientation is deposited to form each of the base section 106, the intermediate section 108, and the top section 110 as depicted in Figure 11 . In the Figure 11 example implementation, after placing the last multi - ply group B, the formation of this portion of the composite panel 100 is complete. Subsequently, additional processing required for a particular process, such as debulking of the prepreg and resin curing, can be completed.

[0057] Premanufacturing the multi-layer laminate reduces the production time required at the mandrel and thus increases the production volume. The deposition of the 0° collimated fiber tape requires the fewest number of passes of the fiber placement head 1104 above the mandrel 1102 and thus requires the least amount of time compared to forming collimated fiber tapes with 45°, 90°, and -45° orientations. However, one or more multi-layer laminates including one or more 0° reinforcement laminates are contemplated and depicted, for example, in multi-layer laminates C-F. Additionally, although Figure 11 ten multi-layer laminates A-J are depicted, each having two or three reinforcement laminates, it should be understood that any number of multi-layer laminates having any number of reinforcement laminates in each group may be contemplated.

[0058] Thus, in a method for forming a composite panel, a plurality of reinforcement laminates (e.g., carbon fiber laminates including collimated fiber tapes containing carbon fibers) may be stacked together to form a carbon fiber stack of the composite panel segment. The carbon fiber stack may include, for example, at least 8 carbon fiber laminates, or at least 10 carbon fiber laminates, or at least 16 carbon fiber laminates, or at least 100 carbon fiber laminates, or at least 200 carbon fiber laminates, or at least 1000 carbon fiber laminates. For example, a process including the following steps is used to bond together the plurality of carbon fiber laminates: compressing the carbon fiber laminates and curing the resin pre-impregnating the carbon fiber laminates. The plurality of carbon fiber laminates are arranged 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 laminate of the carbon plate stack includes a stiffness determined by the orientation of the carbon fibers within each carbon fiber laminate relative to the longitudinal axis. The plurality of carbon fiber laminates of the carbon fiber stack include a plurality of different orientations relative to the longitudinal axis, and thus, the plurality of carbon fiber laminates of the carbon fiber stack include a plurality of different stiffnesses. The different orientations of the plurality of carbon fiber laminates of the carbon fiber stack are self-symmetric or substantially self-symmetric about the midpoint of the carbon fiber stack through the thickness of the composite panel segment, where the midpoint is parallel to the longitudinal axis. When substantially self-symmetric, the different orientations are symmetrically different in no more than two carbon fiber laminates.

[0059] Additionally, the present disclosure includes examples according to the following clauses:

[0060] 1. A composite panel (100), comprising:

[0061] A composite panel segment (104-110) including a plurality of carbon fiber laminates (200-206) arranged along the length and width of the composite panel segment (104-110), wherein the length of the composite panel segment (104-110) includes a longitudinal axis (A);

[0062] The plurality of carbon fiber laminates (200-206) are stacked into a carbon fiber stack that provides at least a portion of the thickness of the composite panel segment (104-110);

[0063] The carbon fiber stack includes at least eight carbon fiber laminae (200 - 206);

[0064] Each carbon fiber lamina of the carbon fiber stack includes a stiffness, wherein the stiffness of each carbon fiber lamina is determined by the orientation of the carbon fibers within each carbon fiber lamina relative to the longitudinal axis (A);

[0065] The plurality of carbon fiber laminae (200 - 206) of the carbon fiber stack include a plurality of different orientations relative to the longitudinal axis (A), and thus, the plurality of carbon fiber laminae (200 - 206) of the carbon fiber stack include a plurality of different stiffnesses; and

[0066] The different orientations of the plurality of carbon fiber laminae (200 - 206) of the carbon fiber stack are self - symmetric or substantially self - symmetric about the mid - points (400, 500, 600) of the carbon fiber stack through the thickness of the composite plate segments (104 - 110), wherein the mid - points (400, 500, 600) are parallel to the longitudinal axis (A), and wherein, when substantially self - symmetric, the different orientations are self - symmetrically different in no more than two carbon fiber laminae (200 - 206).

[0067] 2. The composite plate (100) according to Clause 1, wherein:

[0068] The longitudinal axis (A) of the composite plate segments (104 - 110) is oriented at 0°; and

[0069] The different orientations of the plurality of carbon fiber laminae (200 - 206) include 0°, 45°, 90°, and - 45° relative to the longitudinal axis (A) of the composite plate segments (104 - 110).

[0070] 3. The composite plate (100) according to Clause 2, wherein the different orientations of the plurality of carbon fiber laminae (200 - 206) of the carbon fiber stack are self - symmetric through the thickness of the composite plate segments (104 - 110).

[0071] 4. The composite plate (100) according to Clause 2, wherein:

[0072] The composite plate segments (104 - 110) are intermediate segments (108) located between a base segment (106) and a top segment (110);

[0073] The intermediate segment (108) has an axial elastic modulus ranging from about 11.4 megapounds per square inch (msi) to about 13.4 msi and a Poisson's ratio ranging from about 0.38 to about 0.50; and

[0074] 46% to 58% of the plurality of carbon fiber laminae (200 - 206) are oriented at 0°.

[0075] 5. The composite panel (100) according to Clause 1, wherein the plurality of carbon fiber laminae (200 - 206) of the carbon fiber stack includes a plurality of repeatable groups of carbon fiber laminae (200 - 206), and each repeatable group of carbon fiber laminae (200 - 206) repeats 2, 4, or 6 times.

[0076] 6. The composite panel (100) according to Clause 1, wherein the different orientations of the plurality of carbon fiber laminae (200 - 206) of the carbon fiber stack are substantially self - symmetric through the thickness of the composite panel segments (104 - 110).

[0077] 7. The composite panel (100) according to Clause 1, wherein the composite panel segments (104 - 110) are intermediate segments (108), the carbon fiber stack is an intermediate segment carbon fiber stack, and the composite panel (100) further comprises:

[0078] A base segment (106) and a top segment (110), wherein the intermediate segment (108) is located between the base segment (106) and the top segment (110);

[0079] The base segment (106) comprises:

[0080] A plurality of carbon fiber laminae (202) arranged along the length and width of the base segment (106), and the length of the base segment (106) is parallel to the longitudinal axis (A);

[0081] The plurality of carbon fiber laminae (202) of the base segment (106) are laminated into a base segment carbon fiber stack that provides at least a portion of the thickness of the base segment (106);

[0082] The base segment carbon fiber stack comprises at least 10 carbon fiber laminae (202);

[0083] The plurality of carbon fiber laminae (202) of the base segment carbon fiber stack includes a plurality of different orientations with respect to the longitudinal axis (A), and thus, the plurality of carbon fiber laminae (202) of the base segment carbon fiber stack includes a plurality of different stiffnesses; and

[0084] The different orientations of the plurality of carbon fiber laminae (200 - 206) of the base segment carbon fiber stack are self - symmetric or substantially self - symmetric about the mid - point (600) of the base segment carbon fiber stack through the thickness of the base segment (106), wherein the mid - point (600) of the base segment (106) is parallel to the longitudinal axis (A), and wherein, when substantially self - symmetric, the different orientations of the plurality of carbon fiber laminae (200 - 206) of the base segment carbon fiber stack are self - symmetrically different in no more than two carbon fiber laminae (200 - 206).

[0085] 8. The composite panel (100) according to Clause 7, wherein:

[0086] The longitudinal axis (A) of the composite panel segment (104 - 110) is oriented at 0°; and

[0087] The different orientations of the plurality of carbon fiber laminae (204) of the intermediate segment carbon fiber stack and the plurality of carbon fiber laminae (202) of the base segment carbon fiber stack include 0°, 45°, 90°, and -45° relative to the longitudinal axis (A) of the composite panel segment (104 - 110).

[0088] 9. The composite panel (100) according to clause 8, wherein:

[0089] The intermediate segment (108) has an axial modulus of elasticity ranging from about 11.4 megapounds per square inch (msi) to about 13.4 msi; and

[0090] The base segment (106) and the top segment (110) each have an axial modulus of elasticity ranging from about 10.0 msi to about 11.4 msi.

[0091] 10. The composite panel (100) according to clause 8, wherein:

[0092] The intermediate segment (108) has a Poisson's ratio ranging from about 0.38 to about 0.50; and

[0093] The base segment (106) and the top segment (110) each have a Poisson's ratio ranging from about 0.36 to about 0.50.

[0094] 11. The composite panel (100) according to any one of clauses 7 to 10, wherein the top segment (110) includes:

[0095] A plurality of carbon fiber laminae (206) arranged along the length and width of the top segment (110), and the length of the top segment (110) is parallel to the longitudinal axis (A);

[0096] The plurality of carbon fiber laminae (206) of the top segment (110) are laminated into a top segment carbon fiber stack that provides at least a portion of the thickness of the top segment (110);

[0097] The top segment carbon fiber stack includes at least 10 carbon fiber laminae (206);

[0098] The plurality of carbon fiber laminae (206) of the top segment carbon fiber stack include a plurality of different orientations relative to the longitudinal axis (A), and thus, the plurality of carbon fiber laminae (206) of the top segment carbon fiber stack include a plurality of different stiffnesses;

[0099] The different orientations of the plurality of carbon fiber laminae (206) of the top section carbon fiber stack are self-symmetric or substantially self-symmetric about the midpoint (700) of the top section carbon fiber stack through the thickness of the top section (110), wherein the midpoint (700) of the top section carbon fiber stack is parallel to the longitudinal axis (A), and wherein, when substantially self-symmetric, the different orientations of the plurality of carbon fiber laminae (206) of the top section carbon fiber stack are symmetrically different in no more than two carbon fiber laminae (206); and

[0100] The plurality of carbon fiber laminae (206) of the top section carbon fiber stack are mirror-symmetric with the plurality of carbon fiber laminae (202) of the base section carbon fiber stack.

[0101] 12. The composite panel (100) according to any one of clauses 7 to 10, wherein:

[0102] The top section carbon fiber stack, the middle section carbon fiber stack, and the base section carbon fiber stack are substantially self-symmetric about the midpoint (500) of the middle section carbon fiber stack; and

[0103] The top section carbon fiber stack, the middle section carbon fiber stack, and the base section carbon fiber stack form at least a part of the spar (102) of the composite panel (100).

[0104] 13. The composite panel (100) according to any one of clauses 7 to 10, further comprising:

[0105] A skin section (104), wherein the top section (110), the middle section (108), and the base section (106) are located on one side of the skin section (104), and the base section (106) is located between the middle section (108) and the skin section (104);

[0106] The skin section (104) comprises:

[0107] A plurality of carbon fiber laminae (200) arranged along the length and width of the skin section (104), wherein the length of the skin section (104) is parallel to the longitudinal axis (A);

[0108] The plurality of carbon fiber laminae (200) of the skin section (104) are laminated into a skin section carbon fiber stack that provides at least a part of the thickness of the skin section (104);

[0109] The skin section carbon fiber stack comprises at least 8 carbon fiber laminae (200);

[0110] The plurality of carbon fiber laminae (200) of the skin section carbon fiber stack include a plurality of different orientations with respect to the longitudinal axis (A), and thus, the plurality of carbon fiber laminae (200) of the skin section carbon fiber stack include a plurality of different stiffnesses; and

[0111] The different orientations of the multiple carbon fiber laminae (200) of the skin section carbon fiber stack are self-symmetric about the midpoint (400) of the skin section carbon fiber stack through the thickness of the skin section (104), wherein the midpoint (400) of the skin section carbon fiber stack is parallel to the longitudinal axis (A).

[0112] 14. The composite panel (100) according to clause 13, wherein:

[0113] The longitudinal axis (A) of the composite panel section (104-110) is oriented at 0°;

[0114] The different orientations of the multiple carbon fiber laminae (204) of the intermediate section carbon fiber stack, the multiple carbon fiber laminae (202) of the base section carbon fiber stack, and the multiple carbon fiber laminae (200) of the skin section carbon fiber stack include 0°, 45°, 90°, and -45° relative to the longitudinal axis (A) of the composite panel section (104-110);

[0115] Approximately 46% to approximately 58% of the multiple carbon fiber laminae (204) of the intermediate section carbon fiber stack are oriented at 0°; and

[0116] Approximately 40% to approximately 46% of the multiple carbon fiber laminae (206, 202) of each of the carbon fiber stacks in the top section carbon fiber stack and the base section carbon fiber stack are oriented at 0°.

[0117] 15. The composite panel (100) according to clause 14, wherein:

[0118] The intermediate section (108) has an axial elastic modulus of approximately 11.4 million pounds per square inch (msi) to approximately 13.4 msi and a Poisson's ratio of approximately 0.38 to approximately 0.50;

[0119] The base section (106) and the top section (110) each have an axial elastic modulus of approximately 10.0 msi to approximately 11.4 msi;

[0120] The base section (106) and the top section (110) each have a Poisson's ratio of approximately 0.36 to approximately 0.50; and

[0121] The skin section (104) has an axial elastic modulus of approximately 7.7 msi to approximately 10.0 msi.

[0122] 16. The composite panel (100) according to clause 15, wherein:

[0123] The top section (110), the intermediate section (108), and the base section (106) form at least a part of a solid laminate; and

[0124] The mismatch in Poisson's ratio between the solid laminate and the skin section (104) is in the range of approximately -0.06 to approximately 0.06.

[0125] 17. A composite panel (100), comprising:

[0126] Composite panel sections (104 - 110) including a plurality of reinforcing laminae (200 - 206) arranged along the length and width of the sections (104 - 110), wherein the length of the composite panel sections (104 - 110) includes a longitudinal axis (A);

[0127] The plurality of reinforcing laminae (200 - 206) are laminated into a reinforcing lamina stack that provides at least a portion of the thickness of the composite panel sections (104 - 110);

[0128] The plurality of reinforcing laminae (200 - 206) of the reinforcing lamina stack include a plurality of different orientations relative to the longitudinal axis (A); and

[0129] The different orientations of the plurality of reinforcing laminae (200 - 206) of the reinforcing lamina stack are self - symmetric or substantially self - symmetric about a mid - point (400, 500, 600, 700) of the reinforcing lamina stack through the thickness of the composite panel sections (104 - 110), wherein the mid - point (400, 500, 600, 700) is parallel to the longitudinal axis (A), and wherein when substantially self - symmetric, the different orientations are symmetrically different in no more than two carbon fiber laminae (200 - 206).

[0130] 18. The composite panel (100) according to clause 17, wherein:

[0131] The plurality of reinforcing laminae (200 - 206) include at least 16 reinforcing laminae (200 - 206); and

[0132] The different orientations of the plurality of reinforcing laminae (200 - 206) are self - symmetric.

[0133] 19. A method of forming a composite panel (100), comprising:

[0134] Stacking a plurality of carbon fiber laminae (200 - 206) together to form a carbon fiber stack of the composite panel sections (104 - 110), wherein the carbon fiber stack includes at least 10 carbon fiber laminae (200 - 206); and

[0135] Bonding the plurality of carbon fiber laminae (200 - 206) together, wherein:

[0136] The plurality of carbon fiber laminae (200 - 206) are arranged along the length and width of the composite panel sections (104 - 110);

[0137] Multiple carbon fiber laminae (200 - 206) form at least a portion of the thickness of the composite plate segments (104 - 110);

[0138] The length of the composite plate segments (104 - 110) includes a longitudinal axis (A);

[0139] Each carbon fiber lamina of the carbon plate stack includes a stiffness;

[0140] The stiffness of each carbon fiber lamina is determined by the orientation of the carbon fibers within each carbon fiber lamina relative to the longitudinal axis (A);

[0141] The multiple carbon fiber laminae (200 - 206) of the carbon fiber stack include multiple different orientations relative to the longitudinal axis (A), and thus, the multiple carbon fiber laminae (200 - 206) of the carbon fiber stack include multiple different stiffnesses;

[0142] The different orientations of the multiple carbon fiber laminae (200 - 206) of the carbon fiber stack are self - symmetric or substantially self - symmetric about the mid - points (400, 500, 600, 700) of the carbon fiber stack within the thickness of the composite plate segments (104 - 110);

[0143] The mid - points (400, 500, 600, 700) are parallel to the longitudinal axis (A); and

[0144] When substantially self - symmetric, the different orientations are self - symmetrically different in no more than two carbon fiber laminae (200 - 206).

[0145] 20. The method according to clause 19, further comprising:

[0146] Orienting the longitudinal axis (A) of the composite plate segments (104 - 110) to 0°;

[0147] Orienting each carbon fiber lamina of the multiple carbon fiber laminae (200 - 206) to have an orientation of one of 0°, 45°, 90°, and - 45° relative to the longitudinal axis (A); and

[0148] Stacking at least 10 carbon fiber laminae (200 - 206) such that the different orientations of the multiple carbon fiber laminae (200 - 206) of the carbon fiber stack are self - symmetric.

[0149] While the numerical ranges and parameters setting forth the broad scope of this teaching are approximations, the numerical values set forth in the specific examples are reported as precisely as possible. Any numerical value, however, inherently contains certain errors necessarily resulting from the standard deviation found in their respective testing measurements. Moreover, all ranges disclosed herein are to be understood to encompass any and all subranges subsumed therein. For example, a range of "less than 10" can include any and all subranges between the minimum value of zero and the maximum value of 10 (and inclusive thereof), that is, any and all subranges having a minimum value equal to or greater than zero and a maximum value equal to or less than 10, such as, for example, 1 to 5. In certain cases, the numerical values assigned to the parameters can take negative values. In such cases, the example range values set forth as "less than 10" can assume negative values, such as, for example, -1, -2, -3, -10, -20, -30, etc.

[0150] Although the present teachings have been shown with respect to one or more implementations, changes and / or modifications may be made to the illustrated examples without departing from the spirit and scope of the appended claims. For example, it will be recognized that although a process is described as a series of actions or events, the present teachings are not limited by the order of these actions or events. Some actions may occur in a different order and / or concurrently with other actions or events not described herein. Also, not all process stages may be required to implement a method in accordance with one or more aspects or implementations of the present teachings. It should be understood that structural components and / or process stages may be added, or existing structural components and / or process stages may be removed or modified. Additionally, one or more of the actions depicted herein may be performed in one or more separate actions and / or stages. Further, to the extent that the terms "comprising," "including," "having," "owning," or variants thereof are used in the detailed description and claims, such terms are intended to be inclusive in a manner similar to the term "including." The term "at least one of..." is used to denote that one or more of the listed items may be selected. As used herein, the term "one or more of..." with respect to a list of items such as, for example, A and B, refers to A alone, B alone, or A and B. Further, in the discussion and claims herein, the term "on" as used with respect to two materials, one "on" the other, refers to at least some contact between the materials, while "above" refers to the materials being in the vicinity, but may have one or more additional intermediate materials such that contact is possible, but not required. Neither "on" nor "above" implies any directionality as used herein. The term "conformal" describes a coating material in which the conformal material retains the angles of the underlying material. The term "about" indicates that the listed value may be changed slightly, so long as the change does not result in non - consistency with the illustrated process or structure. Finally, "exemplary" indicates that the description is used as an example and does not imply that it is ideal. Other implementations of the present teachings will be apparent to those skilled in the art from consideration of the specification and practice of the disclosure herein. The specification and examples are to be considered only as exemplary, with the true scope and spirit of the present teachings being indicated by the appended claims.

[0151] The terms of relative position used in this application are defined based on 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" 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," "on the side" (as in "sidewall"), "higher," "lower," "above," "top," and "below" are defined with respect to the conventional plane or working surface on the top surface of the workpiece, regardless of the orientation of the workpiece.

Claims

1. A composite panel (100) comprising: A solid laminated truss structure (102), the solid laminated truss structure (102) comprising composite panel segments (104 - 110), the composite panel segments (104 - 110) comprising a base segment (106), a top segment (110) and an intermediate segment (108), the intermediate segment comprising a plurality of carbon fiber laminae (200 - 206) arranged along the length and width of the composite panel segment (104 - 110), wherein the length of the composite panel segment (104 - 110) comprises a longitudinal axis (A); The plurality of carbon fiber laminae (200 - 206) are laminated into a carbon fiber stack that provides at least a portion of the thickness of the composite panel segment (104 - 110); The carbon fiber stack comprises at least 8 carbon fiber laminae (200 - 206); Each carbon fiber lamina of the carbon fiber stack comprises a stiffness, wherein the stiffness of each carbon fiber lamina is determined by the orientation of the carbon fibers within each carbon fiber lamina relative to the longitudinal axis (A); The plurality of carbon fiber laminae (200 - 206) of the carbon fiber stack comprise a plurality of different orientations relative to the longitudinal axis (A), and thus, the plurality of carbon fiber laminae (200 - 206) of the carbon fiber stack comprise a plurality of different stiffnesses; The different orientations of the plurality of carbon fiber laminae (200 - 206) of the carbon fiber stack are substantially self - symmetric about a mid - point (400, 500, 600) of the carbon fiber stack through the thickness of the composite panel segment (104 - 110), wherein the mid - point (400, 500, 600) is parallel to the longitudinal axis (A), and wherein the different orientations are self - symmetrically different in no more than two carbon fiber laminae (200 - 206); and Wherein the intermediate segment (108) is located between the base segment (106) and the top segment (110).

2. The composite panel (100) according to claim 1, wherein: The longitudinal axis (A) of the composite panel segment (104 - 110) is oriented at 0°; and The different orientations of the plurality of carbon fiber laminae (200 - 206) comprise 0°, 45°, 90° and - 45° relative to the longitudinal axis (A) of the composite panel segment (104 - 110).

3. The composite panel (100) according to claim 2, wherein: The composite panel segment (104 - 110) is an intermediate segment (108) located between the base segment (106) and the top segment (110); The intermediate segment (108) has an axial elastic modulus of 11.4 megapounds per square inch (msi) to 13.4 msi and a Poisson's ratio of 0.38 to 0.50; and 46% to 58% of the plurality of carbon fiber laminae (200 - 206) are oriented at 0°.

4. The composite panel (100) according to claim 1, wherein the plurality of carbon fiber laminae (200 - 206) of the carbon fiber stack includes a plurality of repeatable groups of carbon fiber laminae (200 - 206), and each repeatable group of carbon fiber laminae (200 - 206) repeats 2, 4, or 6 times.

5. The composite panel (100) according to claim 1, wherein the carbon fiber stack is a middle - section carbon fiber stack, and the composite panel (100) further comprises: A plurality of carbon fiber laminae (202) arranged along the length and width of the base section (106), and the length of the base section (106) is parallel to the longitudinal axis (A); The plurality of carbon fiber laminae (202) of the base section (106) are laminated into a base - section carbon fiber stack that provides at least a portion of the thickness of the base section (106); The base - section carbon fiber stack includes at least 10 carbon fiber laminae (202); The plurality of carbon fiber laminae (202) of the base - section carbon fiber stack include a plurality of different orientations relative to the longitudinal axis (A), and thus, the plurality of carbon fiber laminae (202) of the base - section carbon fiber stack include a plurality of different stiffnesses; And The different orientations of the plurality of carbon fiber laminae (200 - 206) of the base - section carbon fiber stack are self - symmetric or substantially self - symmetric about the mid - point (600) of the base - section carbon fiber stack through the thickness of the base section (106), wherein the mid - point (600) of the base section (106) is parallel to the longitudinal axis (A), and wherein, when substantially self - symmetric, the different orientations of the plurality of carbon fiber laminae (200 - 206) of the base - section carbon fiber stack are self - symmetrically different in no more than two carbon fiber laminae (200 - 206).

6. The composite panel (100) according to claim 5, wherein: The longitudinal axis (A) of the composite panel section (104 - 110) is oriented at 0°; The different orientations of the plurality of carbon fiber laminae (204) of the middle - section carbon fiber stack and the plurality of carbon fiber laminae (202) of the base - section carbon fiber stack include 0°, 45°, 90°, and - 45° relative to the longitudinal axis (A) of the composite panel section (104 - 110); The middle section (108) has an axial elastic modulus of 11.4 megapounds per square inch (msi) to 13.4 msi; The base section (106) and the top section (110) each have an axial elastic modulus of 10.0 msi to 11.4 msi; The middle section (108) has a Poisson's ratio of 0.38 to 0.50; and The base section (106) and the top section (110) each have a Poisson's ratio of 0.36 to 0.

50.

7. The composite panel (100) according to claim 5 or 6, wherein the top section (110) includes: A plurality of carbon fiber laminae (206) arranged along the length and width of the top section (110), and the length of the top section (110) is parallel to the longitudinal axis (A); The plurality of carbon fiber laminae (206) of the top section (110) are laminated into a top section carbon fiber stack that provides at least a portion of the thickness of the top section (110); The top section carbon fiber stack includes at least 10 carbon fiber laminae (206); The plurality of carbon fiber laminae (206) of the top section carbon fiber stack include a plurality of different orientations relative to the longitudinal axis (A), and thus, the plurality of carbon fiber laminae (206) of the top section carbon fiber stack include a plurality of different stiffnesses; The different orientations of the plurality of carbon fiber laminae (206) of the top section carbon fiber stack are self-symmetric or substantially self-symmetric about the midpoint of the top section carbon fiber stack through the thickness of the top section (110), wherein the midpoint of the top section carbon fiber stack is parallel to the longitudinal axis (A), and wherein when substantially self-symmetric, the different orientations of the plurality of carbon fiber laminae (206) of the top section carbon fiber stack are self-symmetrically different in no more than two carbon fiber laminae (206); and The plurality of carbon fiber laminae (206) of the top section carbon fiber stack are mirror-symmetric with the plurality of carbon fiber laminae (202) of the base section carbon fiber stack.

8. The composite panel (100) according to claim 5 or 6, wherein: The top section carbon fiber stack, the middle section carbon fiber stack, and the base section carbon fiber stack form at least a portion of a solid laminated spar structure (102) of the composite panel (100).

9. The composite panel (100) according to claim 5 or 6, further comprising: A skin section (104), wherein the top section (110), the middle section (108), and the base section (106) are located on one side of the skin section (104), and the base section (106) is located between the middle section (108) and the skin section (104); The skin section (104) includes: A plurality of carbon fiber laminae (200) arranged along the length and width of the skin section (104), wherein the length of the skin section (104) is parallel to the longitudinal axis (A); The plurality of carbon fiber laminae (200) of the skin section (104) are laminated into a skin section carbon fiber stack that provides at least a portion of the thickness of the skin section (104); The skin section carbon fiber stack includes at least 8 carbon fiber laminae (200); The plurality of carbon fiber laminae (200) of the skin section carbon fiber stack include a plurality of different orientations relative to the longitudinal axis (A), and thus, the plurality of carbon fiber laminae (200) of the skin section carbon fiber stack include a plurality of different stiffnesses; and The different orientations of the plurality of carbon fiber plies (200) of the skin section carbon fiber stack are self-symmetric about the midpoint (400) of the skin section carbon fiber stack through the thickness of the skin section (104), wherein the midpoint (400) of the skin section carbon fiber stack is parallel to the longitudinal axis (A).

10. The composite panel (100) according to claim 9, wherein: The longitudinal axis (A) of the composite panel section (104-110) is oriented at 0°; The different orientations of the plurality of carbon fiber plies (204) of the intermediate section carbon fiber stack, the plurality of carbon fiber plies (202) of the base section carbon fiber stack, and the plurality of carbon fiber plies (200) of the skin section carbon fiber stack include 0°, 45°, 90°, and -45° relative to the longitudinal axis (A) of the composite panel section (104-110); 46% to 58% of the plurality of carbon fiber plies (204) of the intermediate section carbon fiber stack are oriented at 0°; and 40% to 46% of the plurality of carbon fiber plies (206, 202) of each carbon fiber stack in the top section carbon fiber stack and the base section carbon fiber stack are oriented at 0°.

11. The composite panel (100) according to claim 10, wherein: The intermediate section (108) has an axial elastic modulus of 11.4 million pounds per square inch (msi) to 13.4 msi and a Poisson's ratio of 0.38 to 0.50; The base section (106) and the top section (110) each have an axial elastic modulus of 10.0 msi to 11.4 msi; The base section (106) and the top section (110) each have a Poisson's ratio of 0.36 to 0.50; and The skin section (104) has an axial elastic modulus of 7.7 msi to 10.0 msi.

12. The composite panel (100) according to claim 11, wherein: The top section (110), the intermediate section (108), and the base section (106) form at least a part of a solid laminate; and The mismatch in Poisson's ratio between the solid laminate and the skin section (104) is in the range of -0.06 to 0.

06.

13. A method of forming a composite panel of a solid laminated spar structure, comprising: Stacking a plurality of carbon fiber plies together to form a carbon fiber stack of a base composite panel section, an intermediate composite panel section, or a top composite panel section of the solid laminated spar structure, wherein the carbon fiber stack includes at least 10 carbon fiber plies; And Bonding the plurality of carbon fiber plies together, wherein: The plurality of carbon fiber plies are arranged along the length and width of the composite panel section; The plurality of carbon fiber plies form at least a part of the thickness of the composite panel section; The length of the composite panel section includes a longitudinal axis (A); Each carbon fiber ply of the carbon fiber stack includes a stiffness; The stiffness of each carbon fiber ply is determined by the orientation of the carbon fibers within each carbon fiber ply relative to the longitudinal axis (A). The plurality of carbon fiber plies of the carbon fiber stack include a plurality of different orientations relative to the longitudinal axis (A), and thus, the plurality of carbon fiber plies of the carbon fiber stack include a plurality of different stiffnesses; The different orientations of the plurality of carbon fiber plies of the carbon fiber stack are substantially self-symmetric about the midpoint of the carbon fiber stack through the thickness of the composite panel segment; The midpoint is parallel to the longitudinal axis (A); and The different orientations are symmetrically different in no more than two carbon fiber plies.

14. The method according to claim 13, further comprising: Orienting the longitudinal axis (A) of the composite panel segment at 0°; And Orienting each carbon fiber ply of the plurality of carbon fiber plies to have an orientation of one of 0°, 45°, 90°, and -45° relative to the longitudinal axis (A).

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