Abstract:
Aircraft fuselage structures have fuselage bulkheads in which the bulkhead outer caps are integrated with the skin, thereby reducing fastener count and weight. These outer caps and skin are preferably co-cured to form a strong structure. The outer caps can be advantageously constructed as continuous hoops of pultruded elements. The outer cap need not be interrupted by contours or cutouts for stringers, saving weight and reducing complexity. It is contemplated that by integrating the bulkhead outer caps into the skin, a bulkhead can still maintain equivalent stiffness and strength, while saving a significant number of rivets as compared to a comparable design without the outer cap.

Description:
This application claims priority to U.S. Provisional Application Ser. No. 60/979650 filed Oct. 12, 2007 which is incorporated by reference herein in its entirety. 
    
    
     FIELD OF THE INVENTION 
     The field of the invention is the manufacture of composite structures. 
     BACKGROUND 
     The advantages of composite structures in aircraft construction are well known in the aerospace industry. Composites are increasingly used in the primary structure of new aircraft, both as an alternative to and adjunct to metal structures. Carbon fiber composites in which carbon fibers are embedded in an epoxy matrix offer relatively high stiffness and high strength while maintaining a low weight. 
     In one manner of producing large commercial transport aircraft the fuselage structure consists of a thin load bearing skin supported by lateral stringers to eliminate buckling and circumferential bulkheads to transfer shear and retain the fuselage shape. In metal aircraft construction, bulkheads are commonly joined to the skin by the use of rivets. In some large composite aircraft, including the prior art Boeing™ 787, these bulkheads are joined to the skin with an intermediary clip, necessitating the use of two rows of fasteners, the first between the clip and skin, and the second between the clip and bulkhead. While this method is effective at joining the bulkhead to the skin, it results in an excessive quantity of fasteners, and carries with it significant weight, complexity, and cost penalties. 
     Fuselage and wing skins in aircraft are commonly load bearing in both metal and composite construction. The skins provide a large surface area through which to dissipate shear and bending loads. This large thin structure is susceptible to buckling; therefore an internal support structure is required to adequately distribute loads to the skin and to prevent these skin panels from buckling. Major load carrying components such as wing mounts or landing gear are attached to internal frames which then distribute their concentrated loads out to the distributed skin panels. There are several prior art arrangements of the composite skin and supporting internal structure. 
     Relatively smaller composite aircraft such as the Raytheon™ Premier have used a load carrying skin structure which is supported against buckling by a lightweight honeycomb core. The use of a lightweight core material between skins is referred to as sandwich construction.  FIG. 1  shows the nose fuselage section  100  of a Raytheon™ Premier. An outer skin layer  110  and an inner skin layer  120  surround a honeycomb core layer  130 . 
     Larger composite aircraft such as the Boeing™ 787 have used the aforementioned skin and stringer configuration. In this configuration, stringers support the skin and major loads are transferred to the skin through the bulkheads spaced longitudinally through the aircraft.  FIG. 2  depicts a Boeing™ 787 fuselage barrel section  200 , comprising a skin  210  supported by regularly-spaced longitudinal stringers  220  and circumferential bulkheads  230 . Fasteners  240  are used in attaching the bulkhead to the skin. 
     Other structural construction variations are also possible. In a hybrid configuration, bulkheads are used in combination with cored skin sandwich structure which is interrupted or tapered down at each bulkhead interface. In this configuration, the core prevents buckling of the skin and the bulkhead transfers concentrated loads to the skin. 
     Aircraft composite materials must often be cured to obtain the desired properties. Curing usually involves exposing the structure to combinations of one or all of elevated temperature, elevated pressure, or diminished pressure. A composite structure is considered “co-cured” when all the layers or components of the structure are cured together in a single curing stage, even if some of the layers or components were exposed to some type of curing before the step of co-curing. Co-curing can result in very strong bonds between parts and composite layers. In recognition of this, governmental civil aircraft certification agencies including the FAA currently approve of such co-cured structure without additional riveting between the skin and stringers. 
     Composite structures are often built of assemblies of co-cured parts. Such assembly uses secondary bonding. In compliance with current government certification practice, this secondary bonding between parts takes the form of rivets. 
     Recent composite aircraft fuselage or wing construction as found on the prior art Boeing™ 787 uses co-cured outer skin  210  and stringers  220  in the construction of a fuselage section  200 , which avoids the need for a high number of rivets to attach the stringer  220  to the skin  210 . The stringers  220  are co-cured with the skin  210  and continuous on either side of the bulkhead  230 . As a consequence, the circumferential bulkheads  230  must be contoured around the stringers  210  by means of cutouts  232 . These bulkheads  230  (or ribs in the case of wing construction) are numerous in a typical transport aircraft. The assembly of bulkheads  230  from multiple composite parts on the Boeing™ 787 requires secondary bonding and a large number of rivets  240  for each bulkhead  230 . Each bulkhead  230  is riveted to a series of L-shaped clips around its circumference; each clip in turn is riveted to the skin.  FIG. 3  shows another view of a prior art Boeing™ 787 fuselage section  300 . Floor beams  310 , which separate between the passenger section  320  and the cargo hold section  330 , are built separately and attached to each bulkhead  340  using rivets  350  as secondary fasteners. 
     SUMMARY OF THE INVENTION 
     The present invention provides systems, apparatus, and methods by which the outer cap structural element of an aircraft fuselage bulkhead is integrated into the skin of the aircraft fuselage by layering the skin so as to substantially surround the outer cap, creating a thickened area of the skin, and thereby offering improved aircraft fuselage strength and stiffness and reduced aircraft fuselage weight. 
     It is contemplated that the outer cap of the fuselage bulkhead can be advantageously co-cured with the skin, and potentially fabricated using the same technique. The fuselage bulkhead is preferably coupled to the skin to serve as a structural support using secondary bonding, fasteners, or co-curing. 
     In preferred embodiments, the outer cap is not interrupted by cutouts or contours for stringers, saving weight and reducing complexity. Further, the outer cap can be constructed so as to form a continuous hoop about a cross-section of the fuselage providing increased strength. 
     The outer cap can be constructed to be sufficiently thick to serve as a pad-up, useful for locally increasing stiffness and creating a joint at which to couple composite parts together. Alternatively, it is contemplated that additional pad-up layers can be added to the skin further increasing stiffness. In especially preferred embodiments, the outer cap includes a pultrusion or other pultruded material integrated into the fuselage skin. It is contemplated that the outer cap could also comprise co-cured composite layers with fibers substantially in the hoop direction. 
     The fuselage bulkhead can be advantageously structured to have an inner cap substantially parallel to the outer cap. Further, the bulkhead cross-section for a passenger aircraft can be chosen to have upper and lower portions divided by a floor beam, with a floor beam integral with the bulkhead. The skin-integrated outer caps can also be applied selectively; some bulkheads without outer caps can be integrated into the design. By integrating the bulkhead outer cap into the skin, it is contemplated that a bulkhead might use no more than half of a typical number of rivets used in a comparable design without the outer cap and still maintain equivalent stiffness and strength. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWING 
         FIG. 1  is an illustration of the prior art Raytheon™ Premier composite sandwich fuselage construction. 
         FIG. 2  is an illustration of the prior art Boeing™ 787 fuselage barrel section. 
         FIG. 3  is an illustration of the prior art Boeing™ 787 skin and stringer composite fuselage construction with a floor beam separating upper and lower portions. 
         FIG. 4  is an illustration of a preferred composite fuselage construction with bulkhead outer caps integrated into the fuselage skin with an enlarged view of a portion. 
         FIG. 5  is a detail of a preferred composite fuselage construction. 
         FIG. 6  shows a cross-section of a preferred skin construction showing an area of increased thickness and an integrated outer bulkhead cap. 
         FIG. 7  shows a preferred bulkhead frame along with a section cut. 
     
    
    
     DETAILED DESCRIPTION 
     The present invention provides systems, apparatus, and methods by which the outer structural cap material on a bulkhead is co-cured with and made part of the skin, using secondary bonding only once between the bulkhead frame and skin, and providing a continuous outer cap that is not interrupted by contours or cutouts for stringers. This reduces the number of fasteners used to secure the bulkhead to the skin, and creates a lower complexity, lower weight, higher strength bulkhead. 
     The present invention has utility in aircraft fuselages and aircraft fuselage construction as shown in  FIG. 4 . A preferred aircraft fuselage  400  comprises a skin  410 , a first bulkhead  420 , a second bulkhead  430 , and first and second stringers  440 ,  450 . In especially preferred embodiments, the skin  410  has one or more thickened areas  412  advantageously surrounding an outer cap  414  section around at least a portion of the first bulkhead  420 . Thus, in this example, a thickened area  412  of the skin  410  operates as an outer cap  414  of the bulkhead  420  because the thickened area  412  is coincident with, and is coupled to the bulkhead  420  to provide structural support. In this instance the thickened area  412  is also parallel with a largely lateral cross segment  428 . 
     The first bulkhead  420  preferably includes a cutout  460  to allow the stringer  440  to continue from a near side of the bulkhead  420  to a far side of the bulkhead  420 . The outer cap  414 , however, has no cutouts for stringers. The bulkhead  420  can advantageously comprise largely radial web segments  426 , largely lateral cross segments  428 , and oblique segments. The second bulkhead  430  can optionally have an outer cap section integrated into the skin. 
     As used herein, “composite” means engineered materials made from two or more constituent materials. Of special relevance are carbon composites, in which carbon fiber is embedded in a matrix or resin, especially epoxy matrices, thermosetting or thermoplastic resins. Alternate composites are also contemplated including those containing fiberglass, ceramics, and other elements. 
     In  FIG. 5 , an enlarged portion of an especially preferred fuselage construction is shown. A bulkhead  420  comprises a performed frame with a base segment  424  a largely radial web segment  426  and an upper cross segment  428  in the shape of the letter “J”. The upper cross segment  428  comprises a first layer  522  and second layer  524  of composite material. The first layer  522  comprises a series of largely unidirectional fibers  528  embedded in a resin material. The radial segment  426  is made of a laminated composite material and has layers with fibers  526  arranged in oblique orientations. The aircraft skin  410  has an area of increased thickness  412 , which allows sufficient thickness and strength for rivets  530  and smooth transfer of loads from the bulkhead  420  to the skin  410  and the fuselage  400 . Additional material comprising an outer cap  414  is advantageously embedded in the area of increased thickness  412 . This outer cap  414  is attached to the bulkhead  420  by means of rivets  530  and functions as a bulkhead cap, increasing the bulkhead bending stiffness. The bulkhead web segment  426  is contoured around optional stringers  440 ,  450  by means of cutouts  460 ,  462 . Thus,  FIG. 5  provides another example in which a thickened area  412  of the skin  410  operates as an outer cap  414  of the bulkhead  420 . 
     Prior art fuselage constructions have used two sets of fasteners to couple a bulkhead to the skin, one set used to couple an outer cap to a bulkhead frame, and a second set used to couple the bulkhead outer cap to the skin. The present teachings provide for only a single set of fasteners to be used in coupling a bulkhead to the skin, because the cap and skin are integrated and co-cured. This results in a significant reduction in the fastener count or rivet count. Realistically, the reduction in fastener or rivet count contemplated to fall within the 30-50% range. 
     As used herein, a “laminated” object refers to an object made with laminates, and typically comprises multiple layers or plies of composite with fibers in a resin. Individual layers or plies preferably have a plurality or fibers arranged in a predominantly similar orientation. Different layers in a laminate can have fibers at different angles. However, in some cases, a laminate can comprise only a single layer of material. 
     As used herein, a “pultrusion” or “pultruded material” refers to a pultruded composite. Pultruded composites are typically pieces of composite material with largely constant cross-section formed by pulling fibers through a resin in a die, possibly followed by some form of curing. Because the fibers in a pultrusions are usually pulled through a resin and die, pultrusions often have relatively high compressive strength. After the pultrusion process, subsequent processing or milling can change the cross-section. 
       FIG. 6  depicts a cross section of another preferred bulkhead and laminated composite fuselage skin  600  construction, in which a thickened area of the skin operates as an outer cap of the bulkhead. In preferred embodiments, an outer composite skin layer  610  is accompanied by additional skin layers  612 ,  614  with differing fiber orientations. The second skin layer  612  is shown with fibers  630  embedded in a matrix  632 . The fibers  630  in this skin layer  612  are unidirectional and aligned in the hoop direction. 
     A pultruded outer cap  640  is placed over the outer skin layers  610 ,  612 ,  614 . An additional skin layer  616  is advantageously placed on the other side of the pultruded outer cap  640  to surround the outer cap  640  and make it integral with the skin  600 . The pultruded outer cap features largely unidirectional fibers  642  in the hoop direction. It is contemplated that multiple pultrusions or pieces of pultrusions might be incorporated between skin layers  614 ,  616 . A skin pad-up layer  620  is placed medially (as opposed to laterally or externally) relative to the inner skin layer  616  to locally increase the thickness, strength, and stiffness of the laminated composite skin  600  in the vicinity of a bulkhead  650  and rivets  660 . Thus, the skin  600  has an area of increased thickness brought about by the inclusion of a pultruded outer cap  640  and/or skin pad-up layers  620 . Even a single layer, e.g., 0.005 inches thick, can serve as a pad-up layer. But preferably the pad-up comprises at least first, second and third pad-up layers, each having a thickness of at least 0.01 inches. These pad-up layers would usually be added to the skin during a manufacturing process. The term “added to the skin” means that the pad-up layers are bonded and/or co-cured to/with the skin during the manufacturing process. Contemplated manufacturing processes in this instance include hand lay-up and automated fiber placement. 
     The bulkhead  650  is also of laminated composite construction, with first and second layers  652 ,  654  having suitable fiber orientations. The bulkhead  650  is shown in the shape of the letter Z, but is contemplated to be of any suitable cross section, including those resembling the letters J, Z, I, and C. While it is contemplated that the bulkhead  650  could be co-cured with the skin  600 ,  FIG. 6  depicts secondary bonding in the form of rivets  660  between the bulkhead  650  and laminated composite skin  600 . A rivet  650  has a head  662  that is driven into a countersunk hole  618  in the outer skin layer  610  or layers. The rivet  660  has a tail  664  that sits on a titanium support  666  to prevent delamination. The rivet  660  extends through the skin layers  610 ,  612 ,  614 ,  616 ,  620  and outer cap  640 . 
     All suitable layers and layer thicknesses are contemplated for constructing the skin  600  and bulkhead  650 . A skin layer  616  has a thickness  617  that might realistically be 0.005, 0.01, or even 0.05 inches. The skin  600  might realistically comprise 1, 3, 5, 10, 30, 50, 70, 100, or even 150 layers. It is contemplated that the outer cap  640  can be constructed of one or more pultruded sections or unidirectional laminate pieces stacked side-by-side or on top of each other. The outer cap  640  has a total thickness  647  that might be 0.005, 0.01, 0.05, 0.1, 0.25, 0.5, or even 1.0 inches, which total thickness provides the additional “thickened area”, as for example discussed with respect to  FIGS. 4-6 . The skin  600  and the integrated outer cap  640  are co-cured by any suitable process. Contemplated layers can have any suitable thicknesses can drop to zero outer cap thickness if supporting structures permit. Pultruded strips likewise can be of any workable width and thickness combination. Unless the context dictates the contrary, all ranges set forth herein should be interpreted as being inclusive of their endpoints, and open-ended ranges should be interpreted to include only commercially practical values. 
     It is further contemplated that the skin can have a sandwich construction in some areas. As used herein, “sandwich” construction means a lower density and relatively thicker core between two higher-density skin layers. Preferred sandwich constructions include laminated carbon composite skin layers and honeycomb or foam core. 
     In areas where the skin thickness is at the minimum acceptable gauge, including the outboard sections on a wing or the nose and tail of a fuselage, the cap can serve as a pad-up. In areas of heavy gauge skin thickness, such as at the center of the fuselage, there may not be a need for a padup or cap, as sufficient riveting and load transfer material exists in the skin. The outer cap material can be laminated as part of the layup of the fuselage skin using the same manufacturing process, including fiber placement. 
     In especially preferred configurations the inner cap material is a unidirectional pultruded strip which can be tailored to the correct width and thickness allowing for rapid lay-up of the cap material and increased strength of the cap, especially compressive strength. 
     In  FIG. 7 , a preferred complete bulkhead frame  700  is shown having an upper portion  720  and a lower portion  724 , in a “double bubble” configuration. Contoured areas consisting of cutouts  710 ,  712  are provided to allow stringers to pass through without interruption. A cross-section  730  of the bulkhead frame  700  is shown with a radial web segment, a lateral cross segment  732  that functions as an inner cap  722 ,  726 , and a base segment  736  that allows for attachment to the fuselage skin. Fibers  740  in a layer of the lateral cross segment  732  of the upper portion  720  of the bulkhead frame  700  can advantageously be aligned in the hoop direction  742  or at some angle between the longitudinal direction  744  and the hoop direction  742 . Fibers  748  in the radial web segment  734  can be aligned obliquely at an angle between the radial direction  746  and the hoop direction  742 . The upper portion  720  and lower portion  740  are separated by a floor beam  728 . In a typical passenger aircraft construction, passengers would be seated in the part of the fuselage defined by the upper portion  720  and cargo would be placed in the part defined by the lower portion  724 , with the floor beam  728  defining the passenger compartment floor. 
     In preferred embodiments, the bulkhead  700  has upper and lower portions  720 ,  724 , divided by a floor beam  728 , and the floor beam  728  is integral with the bulkhead  700 . As used herein, the term “integral with” means that the integral elements are co-cured. 
     All suitable fuselage cross-section shapes are contemplated including circular, oval, and other shapes. The upper inner cap  722  and lower inner cap  726  are continuous hoops and are uninterrupted by stringers or cutouts. It may be seen that the upper inner cap  722  and lower inner cap  726  are substantially parallel to the outer cap (not shown) which runs around the circumference of the bulkhead frame  700 , preferably in a continuous hoop. In less preferred embodiments the outer and/or inner caps could be interrupted or discontinuous in some manner. 
     It is contemplated that the outer cap of the bulkhead frame  700  can advantageously be integrated into the skin and co-cured as previously described. It is further contemplated that the outer cap that is integrated with the skin can be a continuous and uninterrupted hoop. 
     In preferred embodiments, the outer cap of the bulkhead, primarily of unidirectional material is layed-up and co-cured with the skin of the aircraft. The inner cap consists of one or two continuous loops of unidirectional material for the upper passenger cabin and cargo bay respectively. This produces a bulkhead with integral floor beam structure which has an uninterrupted inner cap. 
     The previously described systems, apparatus, and methods have application beyond aircraft fuselage construction. For example, in wing construction, an outer cap for a wing rib can be integrated into the wing skin structure. All suitable fibers and matrices are contemplated, including all suitable fiber arrangements and orientations. 
     Thus, specific embodiments and methods for producing a reduced fastener count, lighter and less expensive bulkhead and skin construction have been disclosed. 
     It should be apparent, however, to those skilled in the art that many more modifications besides those already described are possible without departing from the inventive concepts herein. The inventive subject matter, therefore, is not to be restricted except in the spirit of the appended claims. Moreover, in interpreting both the specification and the claims, all terms should be interpreted in the broadest possible manner consistent with the context. In particular, the terms “comprises” and “comprising” should be interpreted as referring to elements, components, or steps in a non-exclusive manner, indicating that the referenced elements, components, or steps may be present, or utilized, or combined with other elements, components, or steps that are not expressly referenced. Where the specification claims refers to at least one of something selected from the group consisting of A, B, C . . . and N, the text should be interpreted as requiring only one element from the group, not A plus N, or B plus N, etc.