Abstract:
A fuselage comprising a skin assembly including an outer, laminate skin bonded to an inner, aluminum doubler. The fuselage also includes a support structure comprising a plurality of longitudinal stringer members and a plurality of annular frame members that are attached to, and cooperate to support, the skin assembly. The aluminum doubler provides additional structural support for the fuselage, and in particular, for the outer laminate skin of the skin assembly. The additional structural strength added by the aluminum doubler allows the use of an improved range of fasteners, such as knife-edge, countersink rivets and further allows the use of the laminate layer even in areas with a large number of cutouts, such as the window track of the fuselage. The members of the support structure may interconnected via a plurality of integral flanges, which, when combined with the skin, provide improved structural strength for the entire fuselage.

Description:
FIELD OF THE INVENTION  
         [0001]    The present invention relates to the construction of aerospace vehicle fuselages, and more particularly, to fuselage assemblies that include laminate skins having alternating metal and non-metal panels.  
         BACKGROUND OF THE INVENTION  
         [0002]    The use of laminate panels in aerospace construction is advantageous as the laminate panels typically have a high strength and a relatively low weight. One problem encountered with laminate panels, however, is the limited commercial availability of large laminate panels. U.S. Pat. No. 5,429,326 by Garesche et al. discloses a system for splicing laminate subpanels to make larger laminate panels for use in an aircraft fuselage. As shown in FIGS. 3A and 3B of Garesche et al., a spliced laminate panel  20  includes alternating metal layers  50  and adhesive layers  51 . The metal layers are made of sections separated by spaced splice lines  55 ,  56 ,  57  and  58 . Ostensibly, the spacing between the splice lines improves the strength of the final assembled panel. The aircraft fuselage includes a support structure comprised of longitudinally extending stringers  24  supported by circumferentially extending frame members  22 . The laminate panels are attached to the stringers and frame members so as to form a skin, as shown in FIG. 1 of Garesche et al. The laminate panels are attached to the stringers using rivets  71 ,  72  that transfixes both the stringer and the panel, as shown in FIG. 7 of Garesche et al.  
           [0003]    Although the splicing system disclosed by Garesche et al. has excellent strength characteristics, improvements in the structural strength for aircraft fuselages are always highly desired. U.S. Pat. No. 5,951,800 to Pettit discloses a splice that includes a plurality of splice straps  20  layered over the staggered splice lines so as to provide local reinforcement for the splice joint. As shown in FIGS.  1 - 3  of Garesche et al., the splice straps are solid sheets of metal that overlie the outermost abutting metal sheets of the laminate structure. The splice straps have sufficient width to exceed the staggered offset between all of the breaks within the splice structure. Thus, the splice straps provide further improvement in the structural strength of the splices used to construct the large laminate sheet for an aircraft fuselage.  
           [0004]    Although some types of fasteners can be used with the large laminate panels, as described by Garesch et al., aircraft manufacturers have relied mostly on bonding for attachment of the laminate skin to the underlying frame and stringer assembly of the fuselage. Reliance on bonding over the use of fasteners is most likely due to concerns about compromising the structural strength of the spliced laminate with the insertion of fasteners. Bonding processes are generally problematic due to the need to anodize the metal being bonded and due to uneven process control during application of the adhesive. In addition, there has been a tendency to avoid placing cutouts through the laminate skins, such as for the insertion of windows, that has led to a preference for limited use of the laminate skins on the fuselage. However, limited use of laminate skins results in “mixed joints, which are joints between the laminate skin and the solid metal skin. It is typically difficult to construct such mixed joints due to the different materials of the laminate and solid metal skins.  
           [0005]    Therefore, it would be advantageous to have a system and method that allows greater employment of laminate materials in an aircraft fuselage so as to improve the strength and reduce the weight of the fuselage. In particular, it would be advantageous to have an aircraft fuselage that includes laminate panels used in areas with a large number of cutouts. Further, it would be advantageous if the laminate panels could be connected to the underlying stringers and frame members in such a way as to improve the structural integrity of the finished fuselage.  
         SUMMARY OF THE INVENTION  
         [0006]    The present invention addresses the above needs and achieves other advantages by providing a fuselage comprising a skin assembly including an outer, laminate skin bonded to an inner, aluminum doubler. The fuselage also includes a support structure comprising a plurality of longitudinal stringer members and a plurality of annular frame members that are attached to, and cooperate to support, the skin assembly. Advantageously, the aluminum doubler provides additional structural support for the fuselage, and in particular, for the outer laminate skin of the skin assembly. The additional structural strength added by the aluminum doubler allows the use of an improved range of fasteners, such as knife-edge, countersink rivets and further allows the use of the laminate layer even in areas with a large number of cutouts, such as the window track of the fuselage. In addition, the members of the support structure may be interconnected via a plurality of integral flanges, which, when combined with the skin, provide improved structural strength for the entire fuselage.  
           [0007]    In one embodiment, the present invention includes an assembly combining a collection of individual parts into a low weight but high strength fuselage for an aircraft. The fuselage assembly includes a plurality of longitudinal stringer members, a plurality of annular frame members, a lightweight doubler and a laminate sheet. The longitudinal stringer members are radially spaced from, and extend generally parallel to, the major longitudinal axis of the fuselage. Further, the longitudinal stringer members are spaced circumferentially from each other. Each of the longitudinal stringer members has a stringer wall structure that includes an outer longitudinal surface. The annular frame members are spaced along the longitudinal axis. Each of the frame members includes a frame wall structure having a plurality of outer circumferential surfaces. Each of the outer circumferential surfaces is structurally spliced by the longitudinal stringer members. The lightweight doubler is attached to, and covers, at least a portion of the outer surfaces of the frame and stringer members. The laminate sheet, comprising alternating layers of metal and composite, is disposed over and attached to the lightweight doubler so as to form an outer skin of the fuselage strengthened by the underlying doubler, the frame members and the stringer members.  
           [0008]    The stringer wall structure of each of the longitudinal stringers may include a flange defining the outer longitudinal surface. Also, the frame wall structure of each of the frame members may include a plurality of flanges, each of the flanges defining a respective one of the outer circumferential surfaces. Each of the flanges of the wall structure overlaps a portion of the flange of each of the respective pair of longitudinal stringer members. Preferably, the overlapping flange portions, the lightweight doubler and the laminate sheet are attached together using a fastener. More preferably, the fastener is a knife-edge, countersunk fastener, such as a rivet.  
           [0009]    Optionally, the laminate sheet may be bonded to the lightweight doubler using an adhesive layer, such as a corrosion inhibiting adhesive layer. Preferably, the surfaces of the doubler and the laminate skin are anodized before application of the adhesive layer.  
           [0010]    Preferably, the metal layers of the laminate skin are aluminum layers and the composite layers are a mixture of fiberglass and epoxy. In addition, the doubler is preferably constructed of a lightweight aluminum.  
           [0011]    The present invention has several advantages. The relatively thick and hard aluminum doubler reduces the stresses around the fasteners in the skin assembly. Such a reduction in the fixation stresses allows the use of a wider range of fastener types, such as the knife-edged, countersunk rivets illustrated herein that have excellent durability. Further, the doubler is easily tailored to local loading conditions (unlike most laminate skins) and is an independent, fail-safe member working with the frame and/or stringer. The doubler also allows the laminate skin to have a constant gauge, or thickness, even in areas having cutouts for receiving windows or areas requiring the use of fasteners. A constant gauge skin is more cost-effective than a customized laminate skin requiring increased thickness in areas around fasteners or cutouts. The combined use of the bond layer and the fasteners results in an improvement in fuselage strength and reliability over the use of bonding alone to attach structural members directly to a laminate skin. In addition, the configuration of the stringer members and the frame members provides for continuous load paths along the length of the stringer members and the circumference of the frame members. The result is an overall increase in the strength of the fuselage without a significant increase in weight. Such an increase in the strength of the fuselage provides the option of using smaller stringer and frame members to reduce the weight of the fuselage. 
       
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0012]    Having thus described the invention in general terms, reference will now be made to the accompanying drawings, which are not necessarily drawn to scale, and wherein:  
         [0013]    [0013]FIG. 1 is a perspective view of a portion of an aircraft fuselage of one embodiment of the present invention;  
         [0014]    [0014]FIG. 2A is a perspective view of a laminate skin panel of the aircraft fuselage shown in FIG. 1;  
         [0015]    [0015]FIG. 2B is a perspective view of a lightweight aluminum doubler of the aircraft fuselage in FIG. 1;  
         [0016]    [0016]FIG. 3 is a perspective view of the inside of the fuselage of FIG. 1 showing a plurality of longitudinal stringer members and annular frame members;  
         [0017]    [0017]FIG. 4 is a perspective view of overlapping portions of one of the stringer members and frame members of FIG. 3;  
         [0018]    [0018]FIG. 5 is an enlarged perspective view of the inside of the fuselage of FIG. 1; and  
         [0019]    [0019]FIG. 6 is an enlarged cross-sectional view showing a connection between the laminate panel, the doubler and the stringer of the fuselage of FIG. 1 using a knife-edge, countersink fastener. 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0020]    The present invention now will be described more fully hereinafter with reference to the accompanying drawings, in which preferred embodiments of the invention are shown. This invention may, however, be embodied in many different forms and should not be construed as limited to the embodiments set forth herein; rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the invention to those skilled in the art. Like numbers refer to like elements throughout.  
         [0021]    An aircraft fuselage  10  of the present invention is shown in FIG. 1. The aircraft fuselage includes a plurality of stringer members  11 , a plurality of frame members  12  and a skin assembly  13  having a row of windows  14  housed therein. The skin assembly  13  is attached to the members  11 ,  12  via an array of fasteners  15  that transfix the skin assembly and the members, as shown in FIG. 3. The fuselage  10  of the illustrated embodiment has a tapered, cylindrical shape frequently used in aircraft, but other shapes are also possible and are still considered to be within the scope of the present invention. It should also be noted that the present invention is applicable to fuselage structures for other craft, including other aerospace craft such as a rocket or a launch vehicle, where it is desirable to have a lightweight, strong structure.  
         [0022]    The skin assembly  13  includes individual panels of an outer laminate skin  18  overlaid on matching panels of an inner lightweight aluminum doubler  20  for additional strength, as shown in FIGS. 2A and 2B. Both the laminate skin  18  and the doubler  20  define window openings  19  and  21 , respectively, which are coincident when the laminate skin and doubler are properly assembled. Such an alignment of the window openings allows them to receive and firmly hold the windows  14 , as shown in FIG. 1. The doubler  20  further defines a plurality of weight reducing cutouts  22  that are positioned to be interspersed between the members  11 ,  12  after the skin is assembled into the fuselage  10 , as shown in FIG. 5.  
         [0023]    The stringer members  11  are elongate members extending generally parallel to the axis of the aircraft fuselage  10 , as shown in FIGS. 1 and 3. The frame members  12  are annular hoops, extending circumferentially around, and spaced along, said axis of the fuselage. The stringer members  11  each include a wall structure having a top flange  26  and a bottom flange  27  connected by a web  29 , as shown in FIG. 4. The bottom flange  27  defines an outer longitudinal surface  28  that abuts the skin assembly  13  of the fuselage  10 . In addition, the bottom flange  27  extends laterally outwards from both sides of the web  29  locally at the frame to splice the frame bottom flange  33 . Bottom flange  27  is tailored between frames to react to flight and pressure loads. The top flange  26  of the stringer member extends laterally outwards from one side of the web  29  and is relatively narrow compared to the height of the stringer member.  
         [0024]    The frame members  12  also each include a wall structure having a top flange  32  and a bottom flange  33  connected by a web  36 . The bottom flange  33  extends laterally outwards from both sides of the web  36 , while the top flange  32  extends laterally outwards from one side of the web  36 . The web of each of the frame members  12  defines a plurality of openings  37 , each of the openings corresponding to one of the stringer members  11 . The bottom flange  33  also defines a plurality of flange openings  38 , each of the flange openings corresponding to one of the stringer members  11 . Together, the web openings  37  and the flange openings  38  allow the stringer members to extend through (i.e., be spliced by) the frame members  12 , along the skin assembly  13  of the fuselage  10 . Preferably, the web  36  of each of the frame members  12  has an area of increased thickness  39  around each of the openings  37  and each of the web openings is preferably circular to guard against crack initiation and propagation during loading of the fuselage  10  and to stabilize top flange  32 . The size, shape and material construction of the members  11 ,  12  can be varied to suit the type of fuselage being assembled, and, therefore, the members as depicted herein should not be considered limiting.  
         [0025]    Because the stringer members  11  extend through the frame members  12 , contact of the bottom flange  33  of each of the frame members  12  with the skin assembly  13  occurs at a plurality of circumferentially oriented surfaces  34  defined by the bottom flange, as shown in FIG. 5. Restated, the flange openings  38  interrupt the contact of the bottom surface of the bottom flange  33  with the skin assembly  13  as the bottom flange extends along the inner surface of the skin assembly. Overlapping portions  35  are formed between the bottom flanges at intersections of the stringer members  11  and frame members  12 . These overlapping portions ensure that the members  11 ,  12  form a plurality of continuous circumferential outer surfaces that arrest crack propagation when attached to the aluminum doubler. Further, the bottom flanges  27  of the stringer members  11  are adjacent to the skin assembly  13  at the overlapping portions  35 , allowing the stringers to be in close contact with the skin along the entire fuselage  10 .  
         [0026]    The fasteners  15  are spaced along the length of the bottom flanges  27 ,  33  of each of the stringer and frame members  11 ,  12 . The fasteners  15  transfix the laminate skin  18 , the doubler  20  and both of the bottom flanges  27 ,  33  in the overlapping portions  35 . Therefore, use of the fasteners  15  in the overlapping portions  35  connects the frame members  12  to the stringer members  11 , and both members to the skin assembly  13 . The respective bottom flanges of the stringer and frame members  11 ,  12  are attached directly to the skin assembly  13  by the fasteners  15  in the non-overlapping portions.  
         [0027]    A typical attachment of the bottom flange  27  of one of the stringer members  11  to the skin assembly  13  using one of the fasteners  15  is shown in FIG. 6. In particular, one of the fasteners  15  is depicted in phantom lines and is a rivet that transfixes the laminate skin  18 , the doubler  20  and the bottom flange  27  of the stringer and ends in a flattened end  44  for a secure fixation. The outer laminate skin  18  includes alternating aluminum foil layers  40  and fiberglass epoxy layers  41 . Preferably, the foil layers  40  are 0.010 inch thick 2024-T3 aluminum and the fiberglass epoxy layers  41  are approximately 0.005 inch thick. A bond layer  42  connects the inner one of the foil layers  40  to the lightweight, aluminum doubler  20 . Preferably, the bond layer is an adhesive bond that is approximately 0.005 inch thick. The outer laminate skin  18  may have a different number, or type, of layers, including variations in layer thickness modified to suit the desired application. For instance, additional, thicker layers  40 ,  41  may be needed for a heavier fuselage, or a fuselage that will be subjected to higher pressures. The doubler is tailored in thickness from 0.015 to 0.18 inches thick, depending upon local loads. Although aluminum is the preferred material for the doubler  20 , other relatively lightweight materials could also be used to construct the doubler.  
         [0028]    Assembly of the fuselage  10  preferably includes preparation of the surfaces to be bonded. The outer surfaces of the individual panels of the aluminum doubler  20  are anodized followed by application of a corrosion inhibiting adhesive primer. The inner surfaces of the individual panels of the laminate skin  18  are also anodized and primed. The adhesive bond layer  42  is applied to the primed surfaces of matching panels of the laminate skin  18  and the doubler  20 . The matching panels are adhered together and, if necessary, cured to dry the bond layer.  
         [0029]    After the panels have been prepared, the stringer members  11  and the frame members  12  are positioned in their desired final configuration. The cured panels of the skin assembly  13  are riveted to the outer surfaces  28 ,  34  of the positioned stringer and frame members. Riveting preferably includes driving knife-edged rivets through the layers of the laminate skin  18 , the doubler  20  and one, or if in an overlapping region  35  both, of the bottom flanges  27 ,  33 . The wedge shaped head end of the rivet is countersunk into the laminate skin  18  and the leading, knife-edged end is deformed into the flattened end  44  after it emerges from the bottom flange, as shown in FIG. 6.  
         [0030]    The present invention has several advantages. The relatively thick and hard aluminum doubler  20  reduces the stresses around the fasteners  15  in the skin assembly  13 . Such a reduction in the fixation stresses allows the use of a wider range of fastener types, such as the knife-edged, countersunk rivets illustrated herein that have excellent durability. Further, the doubler  20  is easily tailored to local loading conditions (unlike most laminate skins) and is an independent, fail-safe member. The doubler also allows the laminate skin  18  to have a constant gauge, or thickness, even in areas having cutouts (such as the window openings  19 ) or areas requiring the use of fasteners. A constant gauge skin is more cost-effective than a customized laminate skin requiring increased thickness in areas around fasteners or cutouts. The combined use of the bond layer  42  and the fasteners  15  results in an improvement in fuselage strength and reliability over the use of bonding alone to attach structural members directly to a laminate skin. In addition, the configuration of the stringer members  11  and the frame members  12  provides for continuous load paths along the length of the stringer members and the circumference of the frame members. The result is an overall increase in the strength of the fuselage  10  without a significant increase in weight. Such an increase in the strength of the fuselage provides the option of using smaller stringer and frame members  11 ,  12 , to reduce the weight of the fuselage.  
         [0031]    Many modifications and other embodiments of the invention will come to mind to one skilled in the art to which this invention pertains having the benefit of the teachings presented in the foregoing descriptions and the associated drawings. Therefore, it is to be understood that the invention is not to be limited to the specific embodiments disclosed and that modifications and other embodiments are intended to be included within the scope of the appended claims. Although specific terms are employed herein, they are used in a generic and descriptive sense only and not for purposes of limitation.