Patent Publication Number: US-11383819-B2

Title: Composite vehicle body

Description:
RELATED APPLICATIONS 
     This application is a continuation application of U.S. States application Ser. No. 15/682,818 that was filed on Aug. 22, 2017, which is based on and claims the benefit of priority from U.S. Provisional Application No. 62/417,056 that was filed on Nov. 3, 2016, the contents of all of which are expressly incorporated herein by reference. 
    
    
     TECHNICAL FIELD 
     The present disclosure relates generally to a vehicle body and, more particularly, to a vehicle body made from a composite material. 
     BACKGROUND 
     A vehicle body (e.g., an airplane body, a car body, or a boat body) generally includes an internal skeleton that gives shape to the vehicle, and a skin that overlays the skeleton and provides a smooth outer surface. Modern vehicle bodies are fabricated from a combination of different materials, including composites. For example, the skeleton is typically made of wood, aluminum, or stainless steel, while the skin is typically made of a fiber (e.g., a carbon fiber or fiberglass) embedded within a resin matrix. 
     Pultrusion is a common way to manufacture straight skeletal parts of a vehicle body (e.g., beams, longerons, etc.). During pultrusion manufacturing, individual fiber strands, braids of strands, and/or woven fabrics are pulled from corresponding spools through a resin bath and through a stationary die. The resin is then allowed to cure and harden. Due to the pulling of the fibers prior to curing, some of the fibers may retain a level of tensile stress after curing is complete. This tensile stress can increase a strength of the skeletal part in the direction in which the fibers were pulled. 
     A vacuum-assisted resin transfer molding (VARTM) process is commonly used to fabricate the skin of the vehicle body, after the internal skeleton has already been formed. In a VARTM process, sheets of fibrous material are manually pulled over the internal skeleton and then tacked in place. The tacked material is then manually coated with a liquid matrix (e.g., a thermoset resin or a heated thermoplastic), covered with a vacuum bag to facilitate impregnation of the liquid matrix, and allowed to cure and harden. 
     Although pultrusion manufacturing and VARTM can be an acceptable ways to produce vehicle body parts in some situations, they can also be problematic. In particular, the VARTM-produced skin is often attached to the pultruded skeletal parts and/or reinforced via metallic fasteners (e.g., screws, rivets, and clips). The use of metallic fasteners can drive skeletal design and increase a weight and cost of the vehicle body. In addition, the various vehicle body parts may need to be joined to each other via specially designed hardware, which can also be heavy and costly. Further, electronics (e.g., sensors, heaters, electrical leads, etc.) may need to be added to the vehicle bodies after manufacture, which can further increase the weight, cost, and unreliability. Finally, conventional pultrusion and VARTM manufacturing processes may provide little flexibility in the design and/or use of the vehicle body. 
     The disclosed composite vehicle body is directed to overcoming one or more of the problems set forth above and/or other problems of the prior art. 
     SUMMARY 
     In one aspect, the present disclosure is directed to a vehicle body. The vehicle body may include an internal skeleton forming a wing shape, and a skin formed over the internal skeleton. The skin may include a matrix material, and a plurality of continuous fibers encased within the matrix material. The plurality of continuous fibers may curve from a base end near a fore/aft center of the wing shape outward toward leading and trailing edges of the wing shape at a tip end. 
     In another aspect, the present disclosure is directed to another vehicle body. This vehicle body may include an internal skeleton, and a skin fabricated in-situ over the internal skeleton. The internal skeleton and the skin together may form a wing shape. The internal skeleton may include a plurality of continuous fibers that extend into and form a portion of the skin. 
     In another aspect, the present disclosure is directed to another vehicle body. This vehicle body may include an internal skeleton, and a skin fabricated in-situ over the internal skeleton. The skin may include a matrix material and a plurality of continuous fibers encased within the matrix material. The matrix material may include multiple resins, each deposited at different strategic locations along a length of the plurality of continuous fibers in the skin. 
     In another aspect, the present disclosure is directed to another vehicle body. This vehicle body may include a skin including matrix material and plurality of continuous fibers at least partially coated with the matrix material. The plurality of continuous fibers may include a first plurality of continuous fibers arranged into a first layer that is generally parallel to an outer surface of the vehicle body, a second plurality of continuous fibers arranged into a second layer overlapping the first layer, and a third plurality of continuous fibers extending normal to the outer surface from the first layer into the second layer to interlock the first and second layers. 
     In another aspect, the present disclosure is directed to another vehicle body. This vehicle body may include an internal skeleton having a plurality of adjacent tubular structures that together form a wing shape, and a skin formed from a composite material over the internal skeleton. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a diagrammatic illustration of an exemplary vehicle body; 
         FIG. 2-5  are diagrammatic illustrations of exemplary portions of the vehicle body of  FIG. 1  during manufacture; 
         FIG. 6  is a cross-sectional illustration of an exemplary fiber that may be used to fabricate the vehicle body of  FIG. 1 ; and 
         FIGS. 7-9  are diagrammatic illustrations of additional exemplary portions of the vehicle body of  FIG. 1 . 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  illustrates an exemplary vehicle body (“body”)  10 . In the disclosed embodiment, body  10  is an aircraft body (e.g., an airplane body or a drone body). It is contemplated, however, that body  10  could be of another type (e.g., a car body, a boat body, etc.), if desired. Body  10 , regardless of its configuration and intended use, may include one or more components (e.g., a fuselage  12 , one or more wings  14 , etc.) made from an internal skeleton (e.g., spars, ribs, stringers, bulkheads, trusses, longerons, etc.)  16  covered by an external skin  18 . In some embodiments, the components of body  10  may be fabricated separately and subsequently joined together (e.g., via threaded fastening, riveting, etc.). In other embodiments, the body components may be fabricated together as an integral monolithic structure (e.g., a structure that cannot be disassembled without at least some destruction). 
     As shown in  FIG. 2 , one or more of the components of body  10  may be fabricated via an additive manufacturing process. For example, skeleton  16  may be fabricated from a first additive manufacturing process, while skin  18  may be manufactured from a second and different additive manufacturing process. It is contemplated that both skeleton  16  and skin  18  could be manufactured from the same additive manufacturing process, if desired. 
     The first additive manufacturing process may be a pultrusion and/or extrusion process that creates hollow tubular structures  20  from a composite material (e.g., a material having a matrix M and at least one continuous fiber F). In particular, one or more heads  22  may be coupled to a support  24  (e.g., to a robotic arm) that is capable of moving head(s)  22  in multiple directions during discharge of structures  20 , such that resulting longitudinal axes  26  of structures  20  are three-dimensional. Such a head is disclosed, for example, in U.S. patent application Ser. Nos. 13/975,300 and 15/130,207 and in PCT Application Number 2016042909, all of which are incorporated herein in their entireties by reference. 
     Head(s)  22  may be configured to receive or otherwise contain the matrix material M. The matrix material M may include any type of liquid resin (e.g., a zero volatile organic compound resin) that is curable. Exemplary resins include epoxy resins, polyester resins, cationic epoxies, acrylated epoxies, urethanes, esters, thermoplastics, photopolymers, polyepoxides, thermoset acrylates, thermosets, bismaleimides, silicon, and more. In one embodiment, the pressure of the matrix material M inside of head(s)  22  may be generated by an external device (e.g., an extruder or another type of pump) that is fluidly connected to head(s)  22  via corresponding conduits (not shown). In another embodiment, however, the pressure may be generated completely inside of head(s)  22  by a similar type of device and/or simply be the result of gravity acting on the matrix material M. In some instances, the matrix material M inside head(s)  22  may need to be kept cool and/or dark in order to inhibit premature curing; while in other instances, the matrix material M may need to be kept warm for the same reason. In either situation, head(s)  22  may be specially configured (e.g., insulated, chilled, and/or warmed) to provide for these needs. 
     The matrix material M stored inside head(s)  22  may be used to coat any number of continuous fibers F and, together with the fibers F, make up walls of composite structures  20 . The fibers F may include single strands, a tow or roving of several strands, or a weave of many strands. The strands may include, for example, carbon fibers, vegetable fibers, wood fibers, mineral fibers, glass fibers, metallic wires, SiC Ceramic fibers, basalt fibers, etc. The fibers F may be coated with the matrix material M while the fibers F are inside head(s)  22 , while the fibers F are being passed to head(s)  22 , and/or while the fibers F are discharging from head(s)  22 , as desired. In some embodiments, a filler material (e.g., chopped fibers) may be mixed with the matrix material M before and/or after the matrix material M coats the fibers F. The matrix material, the dry fibers, fibers already coated with the matrix material M, and/or the filler may be transported into head(s)  22  in any manner apparent to one skilled in the art. The matrix-coated fibers F may then pass over a centralized diverter (not shown) located at a mouth of head(s)  22 , where the resin is caused to cure (e.g., from the inside-out, from the outside-in, or both) by way of one or more cure enhancers (e.g., UV lights and/or ultrasonic emitters)  27 . 
     In the example of  FIG. 2 , structures  20  extend in a length direction of wing  14  and makeup at least a portion of skeleton  16 . Each structure  20  may be discharged adjacent another structure  20  and/or overlap a previously discharged structure  20 , and subsequently cured such that the liquid resin within neighboring structures  20  bonds together. Any number of structures  20  may be grouped together and have any trajectory required to generate the desired skeletal shape of wing  14 . 
     In some embodiments, a fill material (e.g., an insulator, a conductor, an optic, a surface finish, etc.) could be deposited inside and/or outside of structures  20  while structures  20  are being formed. For example, a hollow shaft (not shown) could extend through a center of and/or over any of the associated head(s)  22 . A supply of material (e.g., a liquid supply, a foam supply, a solid supply, a gas supply, etc.) could then be connected with an end of the hollow shaft, and the material forced through the hollow shaft and onto particular surfaces (i.e., interior and/or exterior surfaces) of structure  20 . It is contemplated that the same cure enhancer(s)  27  used to cure structure  20  could also be used to cure the fill material, if desired, or that additional dedicated cure enhancer(s) (not shown) could be used for this purpose. The fill materials could allow one or more of structures  20  to function as fuel tanks, fuel passages, electrical conduits, ventilation ducts, etc. 
     The second additive manufacturing process used to fabricate the exemplary wing  14  of  FIG. 2  may also be a pultrusion and/or extrusion process. However, instead of creating hollow tubular structures  20 , the second additive manufacturing process may be used to discharge tracks, ribbons, and/or sheets of composite material over tubular structures  20  (and/or over other features of skeleton  16 ) to thereby fabricate skin  18 . In particular, one or more heads  28  may be coupled to a support  30  (e.g., to an overhead gantry) that is capable of moving head(s)  28  in multiple directions during fabrication of skin  18 , such that resulting contours of skin  18  are three-dimensional. 
     Head  28  may be similar to head  22  and configured to receive or otherwise contain a matrix material M (e.g., the same matrix material M contained within head  22 ). The matrix material M stored inside head(s)  28  may be used to coat any number of separate fibers F, allowing the fibers F to make up centralized reinforcements of the discharging tracks, ribbons, and/or sheets. The fibers F may include single strands, a tow or roving of several strands, or a weave of multiple strands. The strands may include, for example, carbon fibers, vegetable fibers, wood fibers, mineral fibers, glass fibers, metallic wires, etc. The fibers F may be coated with the matrix material M while the fibers F are inside head(s)  28 , while the fibers F are being passed to head(s)  28 , and/or while the fibers F are discharging from head(s)  28 , as desired. The matrix material, the dry fibers, and/or fibers already coated with the matrix material may be transported into head(s)  28  in any manner apparent to one skilled in the art. The matrix-coated fibers F may then pass through one or more circular orifices, rectangular orifices, triangular orifices, or orifices of another curved or polygonal shape, where the fibers F are pressed together and the resin is caused to cure by way of one or more cure enhancers  27 . 
     In another exemplary embodiment shown in  FIG. 3 , only a single additive manufacturing process is being used to fabricate wing  14 . In particular, the second manufacturing process described above is being used to additively build up layers of skeleton  16  (e.g., of spars and/or stringers) with continuous fibers F and matrix material M and to cover skeleton  16  with additively built up layers of skin  18  of the same or different continuous fibers F and matrix material M. The fibers F making up skeleton  16  may continue over an outer surface of skeleton  16  to become part of skin  18 , such that a continuous mechanical connection is formed between skeleton  16  and skin  18  by the continuous fibers F. In this way, the number of fasteners required to connect skin  18  to skeleton  16  may be reduced (if not eliminated). It is contemplated that support  24  and/or support  30  may be used to move any number of heads  28  during fabrication of wing  14  (or during fabrication of any other component of body  10 ). 
     As described above, the first and second additive manufacturing processes can be extrusion or pultrusion processes. For example, extrusion may occur when the liquid resin matrix M and the associated continuous fibers F are pushed from head(s)  22  and/or head(s)  28  during the movement of supports  24  and/or  30 . Pultrusion may occur after a length of resin-coated fibers is connected to an anchor (not shown) and cured, followed by movement of head(s)  22  and/or heads ( 28 ) away from the anchor. The movement of head(s)  22  and/or head(s)  28  away from the anchor causes the fibers F to be pulled from the respective head(s) along with the coating of the matrix material M. 
     In some embodiments, pultrusion may be selectively implemented to generate tension in the fibers F that make up skeleton  16  and/or skin  18  and that remains after curing. In particular, as the fibers F are being pulled from the respective head(s), the fibers F may be caused to stretch. This stretching can create tension within the fibers F. As long as the matrix M surrounding the fibers F cures and hardens while the fibers F are stretched, at least some of this tension remains in the fibers F and functions to increase a strength of the resulting composite structure. 
     Structures fabricated via conventional pultrusion methods may have increased strength in only a single direction (e.g., in the one direction in which fibers were pulled through the corresponding die prior to manual resin impregnation and curing). However, in the disclosed embodiment, the increased strength in the skeleton  16  and/or skin  18  of body  10  (e.g., within wing  14 ) caused by residual tension within the corresponding fibers F may be realized in the axial direction of each the fibers F. And because each fiber F could be pulled in a different direction when being discharged by head(s)  22  and/or  28 , the tension-related strength increase may be realized in multiple (e.g., innumerable) different directions. 
     Structures fabricated via conventional pultrusion methods may have strength increased to only a single level (e.g., to a level proportionate to an amount in which the fibrous cloth was stretched by the pulling machine prior to manual resin impregnation and curing). However, in the disclosed embodiment, because the matrix M surrounding each fiber F may be cured and harden immediately upon discharge, the force pulling on the fiber F may be continuously varied along the length of the fiber F, such that different segments of the same fiber F are stretched by different amounts. Accordingly, the tensile stress induced within each of the different segments of each fiber F may also be different, resulting in a variable strength within the different segments of skeleton  16  and/or skin  18  of body  10 . This may be beneficial in variably loaded areas of body  10  (e.g., at the intersection of wing  14  and fuselage  12 , within a center of wing  14 , at the leading edge of wing  14 , etc.). 
       FIG. 4  illustrates an exemplary way in which the fibers F of skin  18  can be arranged to provide for desired characteristics of wing  14 . In this example, the fibers F are arranged organically (e.g., in the way that a tree grows or in the way that blood veins are situated in the body). Specifically, the fibers F placed over structure  16  of wing  14  may be anchored at an intersection with fuselage  12  and in a general fore/aft center. The fibers F may then be pulled toward a distal tip of wing  14 , and away from the fore/aft center (e.g., toward a leading or trailing edge of wing  14 ), with different fibers F extending different distances toward the distal tip. In addition, because the discharging matrix material M may cure immediately upon discharge from head  28  and bond to either structure  16  or previously discharged layers of skin  18 , movement of head  28  during discharge may be controlled to create trajectories of the fibers F that curve. It is contemplated that the fibers F may pass completely around wing  14  at its distal termination point, and then be pulled back toward fuselage  12  following a mirror image of its initial trajectory. This arrangement of organically arranged fibers may be located at a top side of wing  14 , at a lower side, and/or around a cross-sectional perimeter of wing  14  at multiple locations. With this arrangement, a greater density of fibers F may exist near fuselage  12  than near the distal tip of wing  14 . Accordingly, wing  14  may be thicker near fuselage  12  and near the general fore/aft center, and taper toward the distal tip and the leading and trailing edges. Other arrangements and/or fiber distribution schemes may be employed, as desired. 
     In one exemplary embodiment shown in  FIG. 5 , some of the fibers F within the composite material making up one or more portions of body  10  have unique characteristics. For example, while a majority of wing  14  may comprise a structural type fiber F s  (e.g., carbon fibers, fiberglass, or Kevlar fibers), some portions of wing  14  may include another type of fiber F (e.g., electrically conductive fibers F ec , optical fibers F o , shape memory fibers F sm , etc.). The other type of fibers F may be selectively interwoven with the structural type fibers F s  at strategic locations. For example, electrically conductive fibers F ec  may be located at leading edges and/or thinner portions of wing  14  and used as heating electrodes that can be connected to a power source and used to remove ice from wing  14 . Alternative, electrically conductive fibers F ec  may be located at high-stress regions (e.g., at the intersection of wing  14  and fuselage  12 ) and used as strain gauges to detect loading of body  10 . In a similar manner optical fibers F o  may be located at high-stress regions and an energy beam passed therethrough. As body  10  flexes, the optical fibers F o  may be squeezed and/or closed, thereby generating an optical feedback signal indicative of the flexing. In yet another embodiment, fibers F sm  fabricated from a shape memory alloy (e.g., Nitonol) may be interwoven with the structural type fibers F s  and selectively energized (e.g., via electricity or heat) to cause flexing (e.g., controlled pulling and/or pushing) of body  10  that results in a desired aerodynamic performance (e.g., steering, orientation, elevation control, stability, drag, etc.). As shown in  FIG. 6 , it is contemplated that the electrically conductive fibers F ec , the optical fibers F o , and/or the shape memory fibers F sm  may be coated with another material (e.g., insulation, a strength enhancing layer, etc.), if desired. It is also contemplated that other electrical components (e.g., resistors, capacitors, etc.) may be extruded through heads  22 ,  28  and/or automatically picked-and-placed (e.g., via attachments associated with heads  22  and/or  28 ) during discharge of fibers F ec , fibers F o , and/or fibers F sm . Operation of these components and/or of fibers F ec , fibers F o , and/or fibers F sm  may be selectively tuned in these instances, for example by adjusting a shape, tension, and/or size of the fibers. 
     Structures fabricated via conventional pultrusion and/or extrusion methods may be limited in the orientation of the associated fibers. That is, the fibers may be generally overlapping and lie in parallel layers. However, in the embodiment illustrated in  FIG. 5 , because the matrix M surrounding each fiber F may be cured and harden immediately upon discharge, the fibers F may be caused to extend into free space without additional support. That is, the fibers F may not be required to lie in flat layers on top of each other. Accordingly, the fibers F making up skeleton  16  and/or skin  18  may be oriented in directions that are perpendicular to each other in three dimensions. For example,  FIG. 5  illustrates fibers F n  that extend in a direction normal to the surface of wing  14 . This may allow for interlocking of fiber layers and/or for the creation of unique (e.g., turbulence enhancing) surface textures. 
     As described above and shown in  FIG. 5 , body  10  may be fabricated as an integral monolithic structure, in some embodiments. For example, wings  14  may be fabricated together with (e.g., at the same time as and without separation from) fuselage  12 . In particular, as support(s)  24  and/or  30  move any number of head(s)  28  over skeleton  16  to create skin  18  (referring to  FIG. 2 ), head(s)  28  may pass from one wing  14 , over or under fuselage  12 , and continue across the opposing wing  14 . In this instance, the fibers F discharging from head(s)  28  may be continuous over wings  14  and fuselage  12 . This process may be repeated any number of times, such that millions (if not hundreds of millions) of fibers F extend through the intersection between wings  14  and fuselage  12 , thereby creating a strong mechanical connection without requiring the use of specialized hardware and/or heavy fasteners. 
     In the exemplary embodiment shown in  FIG. 5 , the matrix M within the composite material making up one or more portions of body  10  has unique characteristics. For example, while a majority of wing  14  may comprise a structural type matrix M s  (e.g., a conventional UV curable liquid resin such as an acrylated epoxy), some portions of wing  14  may include another type of matrix M (e.g., a pyrolized matrix M p , a matrix that remains somewhat flexible M f , etc.). The other type of matrix M may be selectively used to coat the fibers F at strategic locations. For example, the pyrolized matrix M p  may be fed into head  28  as head  28  nears the leading edge of wing  14  and/or a nose of fuselage  12 , such that the resulting composite material may function as a heat shield in these areas. In another example, the flexible matrix M f  may be fed into head  28  as head  28  nears the trailing edge of wing  14  (e.g., where the shape memory fibers F sm  are placed, such that the resulting composite material may be more flexible and selectively warped or twisted to provide for the desired aerodynamic properties described above. 
       FIG. 7  illustrates an exemplary part  32  of skeleton  16  that may be fabricated through the use of head  28  and support  30 . Although depicted and described as a rib of wing  14 , part  32  could be another skeletal component of wing  14  and/or fuselage  12 . In this example, part  32  includes opposing outer support surfaces  34 , opposing internal braces  36 , and a plurality of cross-pieces  38  that interconnect support surfaces  34  and/or braces  36 . It should be noted that outer support surfaces  34  and/or internal braces  36  could have any desired shape, for example curved, flat, stepped, etc. It is also contemplated that outer support surfaces  34  and internal support surfaces  34  could form one or more continuous surfaces, if desired. For example, one or more of support surfaces  34  could be curved and generally tangential with one or more of braces  36  (e.g., at a leading and/or trailing end of the rib). And although cross-pieces  38  are shown as generally straight and oriented at about 45° relative to support surfaces  34  and braces  36 , it is contemplated that cross-pieces  38  could also be curved and/or oriented at another angle. It should be noted that, although seven adjacent and nearly identical parts  32  are shown to make up the disclosed rib, any number of the same or different parts  32  (e.g., only one part  32 ) may be used for this purpose. 
     Part  32  may be created following a unique tool path that allows for use of continuous fibers and provides for high-strength in a low-weight configuration. In particular, part  32  may be fabricated using a middle-out strategy.  FIG. 8  illustrates use of this strategy during fabrication using multiple different and overlapping layers. 
     For example, in a first layer, head  28  may be controlled to start discharging and curing one or more continuous resin-coated fibers at a lower-left corner (e.g., adjacent an internal intersection of a lower support surface  34  and a left brace  36 ), and continue discharging and curing the same resin-coated fiber(s) during travel upward to an adjacent upper-left corner. Head  28  may then move diagonally inward toward a general center of part  32 , and then double back prior to reaching the center to move toward the upper-left corner following a generally parallel trajectory. During this doubling-back maneuver, head  28  may be spaced apart a distance from its original trajectory (e.g., spaced more toward the right of part  32 ), such that an empty space will exist along a diagonal of part  32  and a box shape is formed at internal ends of the diagonal parallel tracks. Head  28  may then move rightward to an upper-right corner of part  32 , followed by about a 90° turn downward upon reaching an internal edge of part  32 . The same general pattern may be repeated at the lower-right corner of part  32  that was made at the upper-left corner, such that a mirror image across a virtual diagonal dividing line is created. Head  28  may then move leftward and stop short of its starting point, after which head  28  may turn through about 45° clockwise and travel diagonally completely across part  32  to the upper-right corner. Head  28  may then double back toward the lower-left corner along a spaced-apart parallel track, such that head  28  is near its starting point (e.g., radially outward and slightly lower than the starting point). During this doubling-back maneuver, head  28  may be spaced apart a distance from its original trajectory (e.g., spaced more toward the left of part  32 ), such that an empty space will exist along a diagonal of part  32 . As head  28  moves towards the upper-right corner, it may deviate from its trajectory at a turn-around point and head into the corner, such that an arrow-head shape is formed at internal ends of the parallel tracks. The arrow-head shape may bond to the hardened fibers previously laid down at this corner location. The diagonally laid fiber(s) may bond to the box shape previously laid down at the center of part  32 . The entire process may be repeated any number of times to add a corresponding number of material tracks to the first layer and thereby increase a cross-sectional area of the first layer. During repetition, part  32  may grow outward and the empty spaces described above as being located between the parallel tracks may be filled in. It should be noted that, during formation of any one layer, the fibers discharging from head  28  may not overlap other fibers such that all fibers are laid down within the same plane. When head  28  reaches an endpoint of a particular layer, the fiber(s) may be cut from head  28 , such that head  28  may be repositioned for start of a new layer. 
     A second layer may be formed directly on top of the first layer, for example by rotating the pattern of the first layer through a desired angle (e.g., through about 90°). By rotating the pattern through about 90°, the fibers extending diagonally completely across part  32  in the second layer may overlap the fibers that doubled back at the center of part  32  in the first layer. This overlapping of different portions of the repeating pattern may help to increase a strength of part  32 . It is contemplated that any number of fibers may be deposited at any location and oriented generally normal to the overlapping layers (e.g., fibers F n  like those shown in  FIG. 5 ) to interlock the layers, if desired. Additionally or alternatively, discharged but uncured portions of a previous layer could be wrapped over subsequently formed layers and then cured to improve an interlock strength. 
     Any number of additional layers may be formed on top of the first two layers in alternating orientations and/or in orientations of incremental rotating angles (e.g., when the angle is not a multiple of 90°). This may continue until a desired thickness of part  32  is achieved. In one example, an entire fuselage  12  and/or wing  14  could be fabricated in this manner. For example, skin  18  could be simultaneously fabricated over part  32  when using the middle-out approach. In particular, an empty space may be created inside of fuselage  12  and/or wing  14  and between adjacent parts  32 , by only creating outer portions of supports  34  and/or braces  36 . 
       FIG. 9  illustrates an alternative method for fabricating skeleton  16  and/or skin  18  of wing  14 . In this embodiment, instead of using the middle-out strategy of  FIG. 8 , different serpentine-patterned layers may be used in alternating and overlapping fashion. For example, when fabricating skeleton  16  of wing  14 , head  28  (referring to  FIG. 7 ) may be controlled to start discharging and curing one or more continuous resin-coated fibers at a right-most or trailing end of a rib cross-section, and continue discharging and curing the same resin-coated fiber(s) during arcuate travel upward and to left. Head  28  may move completely around a top support surface  34  and a leading end of the rib cross-section, and then move back to the starting point in a generally straight trajectory across a lower support surface  34  of the rib cross-section. Upon making a complete circumference of the rib cross-section, head  28  may double back and follow a reverse trajectory during a second lap around the perimeter of the rib. It should be noted that the second lap may be completed radially inside of or outside of the first lap, as desired, as long as both laps remain within the same general plane and immediately adjacent (e.g., bonded to) each other. Although Layer- 1  shown in  FIG. 9  includes two complete laps of material, it is contemplated that only a single lap or more than two laps may be used. During fabrication of the first and/or second laps of Layer- 1 , head  28  may not pass through a center of the rib cross-section, such that the center remains free of cross-pieces and bracing. That is, only a perimeter of the rib cross-section may be formed by Layer- 1 . 
     After completion of Layer- 1 , the fibers trailing from head  28  may be severed, and head  28  may move toward the leading end of the rib cross-section for fabrication of Layer- 2 . Head  28  may be controlled to start discharging and curing one or more continuous resin-coated fibers at top surface  34  of the leading end, and continue discharging and curing the same resin-coated fiber(s) during arcuate travel upward and to right (e.g., toward the trailing end). After forming top support surface  34  and reaching the trailing end of the rib cross-section, head  28  may move back toward the leading end in a generally straight trajectory along lower support surface  34  and then curve around the leading end to the starting point of the first lap of Layer- 2 . During this Layer- 2 , first lap of head  28  around the perimeter of the rib cross-section, head  28  may not pass through the center of the rib, such that the center still remains free of cross-pieces and bracing. 
     Head  28  may pass through the center of the rib cross-section during the second lap of Layer- 2 , such that cross-pieces  38  are created in free-space between opposing nodes in top and lower support surfaces  34 . In particular, after returning to the starting point of the first lap, head  28  may double back and follow the same general trajectory around the curved leading end at a radial location inside of the first lap. After moving around the curved leading end, head  28  may move away from the first lap and angle upward and to the right (i.e., toward the trailing end), passing through the center of the rib cross-section and toward the starting point. In the disclosed embodiment, the angle of cross-piece  38  relative to top and/or lower support surfaces  34  may be about 25-65°. After the material discharging from head  28  engages an inner side of the first lap, head  28  may generally follow the trajectory of the first lap for a short distance, such that a segment of material is laid down adjacent (e.g., at the inside) of the first lap at top support surface  34  of the rib cross-section. In one embodiment, the segment is about ⅓ to ⅙ of the length of the rib cross-section. Thereafter, head  28  may move away from the first lap and angle downward and to the right, again passing through the center of the rib cross-section. After the material discharging from head  28  again engages an inner side of the first lap, head  28  may generally follow the trajectory of the first lap for a short distance, such that another segment of material is laid down adjacent (e.g., at the inside) of the first lap at lower support surface  34  of the rib cross-section. This serpentine pattern of passing through the rib-center and segment-creation may continue until head  28  reaches the trailing end of the rib cross-section. It should be noted that, in the embodiment of  FIG. 9 , cross-pieces  38  within any one the layer of the rib cross-section are only one-sided. That is, a complete “X” through the center of the rib cross-section may be created only by two different layers. 
     Odd layers of the rib cross-section may all be substantially identical to each other, while even layers may be identical to or iterations of each other. For example, Layers- 1 ,  3 ,  5 ,  7 , etc. may all be identical. Likewise, Layers- 2  and  8 ; Layers- 4  and  10 ; and Layers- 6  and  12  may be identical pairs. However, Layers- 2 ,  4 , and  6  may be different iterations of each other. Specifically, the locations of cross-pieces  38  and the top and lower support segments may alternate between the even layers, such that cross-pieces  38  overlap each other and thereby form complete “Xs” and the top and lower support segments line up sequentially to form a complete perimeter around the rib cross-section. It should be noted that, while Layers- 1  through  6  are shown in  FIG. 9  as forming a complete set that can be repeated any number of times during formation of wing  14 , any number of layers may form a set that can be repeated to form wing  14 . 
     In some embodiments, after formation of Layers- 1  through  6  (and/or after multiples of these layers), wing  14  may be complete. That is, the inner lap of each layer may form skeleton  16 , while the outer layer may form skin  18 . In other embodiments, however, a separate skin  18  may be laid over skeleton  16  after formation of Layers- 1  through  6 , if desired. In this embodiment, skin  18  may be laid by hand or by the disclosed system(s). If laid by the disclosed system(s), the second additive process described above (but using sheets of pre-fabricated material instead of individual fibers or tows) may be implemented. 
     INDUSTRIAL APPLICABILITY 
     The disclosed arrangements and designs of skeleton  16  and skin  18  may be used in connection with any type of vehicle body  10 . For example, skeleton  16  and skin  18  may be used in connection with an airplane body, a drone body, a car body, a boat body, or any other type of vehicle body where light-weight, low-cost, and high-performance are important. Vehicle body  10  may be light-weight and low-cost due to the reduction in the number of fasteners required to secure skin  18  to skeleton  16  and/or to secure components of vehicle body  10  to each other. In addition, vehicle body  10  may be light-weight do to the use of composite materials used to make both of skeleton  16  and skin  18 . The high-performance may be provided in the unique ways that particular fibers and resins are used and laid out within skeleton  16  and skin  18 . 
     It will be apparent to those skilled in the art that various modifications and variations can be made to the disclosed vehicle body. Other embodiments will be apparent to those skilled in the art from consideration of the specification and practice of the disclosed vehicle body. It is intended that the specification and examples be considered as exemplary only, with a true scope being indicated by the following claims and their equivalents.