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Embodiments of integral composite panels and joints for composite structures are described In one implementation, an integrated panel spanning substantially the entire wingspan of an aircraft, includes at least a center portion and a pair of outwardly projecting wing portions. The portions may inclu

Embodiments of integral composite panels and joints for composite structures are described In one implementation, an integrated panel spanning substantially the entire wingspan of an aircraft, includes at least a center portion and a pair of outwardly projecting wing portions. The portions may include a skin formed from successive layers or plies of composite material which overlap and offset at the joint between respective sections creating a pad-up area to carry loads between the portions. In a particular implementation, the skin is laid over one or more structural stringers which are transitioned into the joints between sections such as by tapering of the thickness and/or stiffness of the stringer.

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What is claimed is: 1. A composite wing panel, comprising: a frame including a first portion and a second portion that are joined at an interface; composite skin covering the frame, the skin including a plurality of plies of composite material, at least a portion of the plurality of plies of the fi

What is claimed is: 1. A composite wing panel, comprising: a frame including a first portion and a second portion that are joined at an interface; composite skin covering the frame, the skin including a plurality of plies of composite material, at least a portion of the plurality of plies of the first and second portions arranged in an overlapping pattern across an interface between the portions, thereby integrally forming a laminate skin over the interface; wherein the frame includes a plurality of stringers that transition into the interface, the stringers tapered in at least one of height and stiffness as they approach the interface. 2. A composite wing panel according to claim 1, wherein the overlapping pattern includes a non-symmetrical pad-up in the thickness of the skin proximate the interface between the portions, the pad-up being configured to transfer loads across the interface. 3. A composite wing panel according to claim 1 wherein, the plurality of plies from each portion includes a plurality of near zero degree plies arranged lengthwise along a zero degree axis of the respective portion and the overlapping pattern includes overlapping of the near zero degree plies of the first portion with the near zero degree plies of the second portion. 4. A composite wing panel according to claim 1 wherein the interface includes a side-of-body rib; and a plurality of reduced-size bolts for attaching a tapered portion of one of the tapered stringers to the rib. 5. A composite wing panel according to claim 1, wherein the frame includes a center portion, a first wing portion, and a second wing portion, and wherein: the center portion is configured to extend across an aircraft fuselage such that a zero degree axis of the center portion is substantially transverse to a longitudinal axis of the fuselage; the first and second wing portions are integrally formed to the center portion on opposing sides of the center portion and are swept with respect to the center portion such that a zero degree axis corresponding to each wing portion is arranged at a relative sweep angle with respect to the zero degree axis of the center portion; each of the first and second wing portions and the center portion includes a plurality of near zero degree plies arranged approximately lengthwise along the zero degree axis of the respective portion; and the overlapping pattern includes overlapping of the near zero degree plies of the center portion with the near zero degree plies of each wing portion, such that one or more pad-ups in the thickness of the skin are produced between the center portion and each wing portion. 6. The composite wing panel according to claim 5, wherein the overlapping of the near zero degree plies of the wing portions with the near zero degree plies of the center portion occurs at the respective sweep angle such that the near zero degree plies of each wing portion, when subject to a load, act as shear plies in the center section to transfer at least a portion of the load to the center section via the angular shear between the overlapping plies. 7. A composite wing panel according to claim 1, wherein the stringers are tapered in stiffness as they approach the interface. 8. The composite wing panel according to claim 1, wherein endpoints of the plies from each portion are offset from one from another such that the pad-up occurs gradually. 9. The composite wing panel according to claim 1, wherein the stringers transition from a region of relatively high thickness to a region of relatively lower thickness proximate to the interface between the portions. 10. An aircraft, comprising: a fuselage; a composite wing structure joined to the fuselage, the wing structure including a frame that forms a center portion spanning the fuselage and a plurality of wing portions extending laterally outward from the fuselage; wherein the plurality of wing portions are integrally formed to the center portion via a plurality of interleaved composite plies; and wherein the frame includes a plurality of stringers that transition into interfaces of the portions, the stringers tapered in at least one of height and stiffness as they approach the interfaces. 11. The aircraft according to claim 10 wherein at least some of the plurality of plies are formed from one or more strips of carbon fiber tape extending approximately lengthwise along each portion; the overlapping pattern includes an alternating pattern such that plies from the first portion alternate with plies from second portion across the interface; the overlapping pattern includes a pad-up in a thickness of the skin at the interface between the portions; the endpoints of the plies from each portion being offset one from another such that the pad-up occurs gradually; and the pad-up is configured to be positioned proximate an outer surface of a fuselage when the integrally formed portions are coupled to the fuselage, the pad up being configured to be coupled to the fuselage and to transfer loads across the interface between the portions. 12. The aircraft according to claim 10, wherein: the plurality of interleaved plies includes a plurality of plies arranged approximately axially along a zero degree axis associated with each of the portions; and the interleaving includes: overlapping of the zero degree plies of each of the wing portions with the zero degree plies of the center portion to form an area of increased thickness proximate to an intersection of each wing portion with the center portion; and offsetting the endpoints of the overlapping plies of each portion in the area of increased thickness to build the thickness gradually. 13. The aircraft according to claim 12, wherein the offset of the endpoints is within a range of about 0.25 inches to about 1.25 inches per ply. 14. The aircraft according to claim 12, wherein the interfaces include side-of-body joints joining the composite wing structure to the fuselage, wherein: each intersection of the center portion with one of the plurality of wing portions occurs substantially along a body line of the fuselage on a respective side of the aircraft, such that the corresponding area of increased thickness formed extends along the body line; each of the side-of-body joints being proximate to a respective one of the intersections and coupled to the composite wing structure through the corresponding area of increased thickness; and the areas of increased thickness being configured to bear loads transferred between the plurality of wing portions and the center portion. 15. The aircraft according to claim 12, wherein: each of said wing portions comprise swept wing portions, each swept wing portion including a plurality of stringers extending lengthwise substantially parallel to the zero degree axis of the respective portion; the plurality of stringers are configured to receive and support the corresponding plurality of interleaved composite plies; and at least one said stringer corresponding to one said portion is configured to have tapering stiffness along the length of the stringer such that the stringer transitions from a region of relatively high stiffness to a region of relatively lower stiffness approaching an end of the stringer located proximate to one said area of increased thickness. 16. A composite panel, comprising: a first section and a second section coupled to the first section at an interface via a plurality of composite plies arranged axially along a zero degree axis of each first and second section, the plurality of plies of the first section being interleaved with the plurality of plies of the second section; and a framework coupled to the first and second sections and having a plurality of elongated stringers arranged to run lengthwise along each said section, one or more of the elongated stringers from at least one said section being configured to taper as the stringer approaches the interface. 17. The composite panel as recited in claim 16, wherein at least one of the first and second sections is formed from successive layers of a fiber-reinforced composite tape. 18. The composite panel as recited in claim 16, wherein the tapering comprises a reduction of thickness as the stringer approaches the interface between the first and second sections. 19. The composite panel as recited in claim 16, wherein the tapering comprises a reduction of the stiffness of the stringer as the stringer approaches the interface between the first and second sections. 20. The composite panel as recited in claim 19, wherein the stiffness is varied between about 16 MSI and about 2 MSI.