Abstract:
A composite structure comprises stacked sets of laminated fiber reinforced resin plies and metal sheets. Edges of the resin plies and metal sheets are interleaved to form a composite-to-metal joint connecting the resin plies with the metal sheets.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is a continuation-in-part of U.S. patent application Ser. No. 12/857,835 filed Aug. 17, 2010, the entire disclosure of which is incorporated by reference herein. 
     
    
     BACKGROUND INFORMATION 
       [0002]    1. Field 
         [0003]    This disclosure generally relates to composite structures, especially fiber reinforced resin laminates, and deals more particularly with a hybrid composite having a composite-to-metal joint, as well as to a bonded metal laminate used in the joint. 
         [0004]    2. Background 
         [0005]    Bonding techniques are often used to assemble composite structures. In applications where the composite structure also requires fasteners, the local thickness or gauge of the structure surrounding the fastener may need to be increased in order to withstand loads transmitted through the fastener joint. As the local thickness of the structure increases, the fastener may need to be lengthened, thereby adding weight to the structure. Additionally, the increased local thickness of the structure may increase the eccentricity of the load path across the fastener joint, which may place undesired bending loads on the fastener. 
         [0006]    One solution to the problems mentioned above consists of attaching metal fittings to the composite structure in the area of the fasteners. These metal fittings may be formed of titanium or similar metals that may not substantially chemically react with carbon fiber reinforced composites in which they are in contact. Titanium fittings, however may be relatively expensive, particularly when it is necessary to form them into complex shapes. 
         [0007]    Accordingly, there is a need for a composite resin-to-metal joint that may be used to connect substantially all metal fittings with substantially all composite resin structures, which is relatively inexpensive and easy to manufacture, and which may withstand loads transferred around fastener connection points. There is also a need for a composite resin-to-metal joint that substantially avoids chemical reactions between the all metal fitting and the all composite resin structure. Also, there is a need for a composite-to-metal joint that may reduce residual stresses in the joint following a thermal curing. Further there is a need for a bonded metal laminate that may be used in the joints and in other applications where additional strength and durability are required. 
       SUMMARY 
       [0008]    The disclosed embodiments provide a hybrid-type composite structure that includes a fiber reinforced resin composite-to-metal joint that may be used to connect a substantially all-metal fitting with a substantially all composite resin structure or a different structure. The joint provides a transition between the composite and metallic structures that is suitable for use in higher performance applications, such as aerospace vehicles. This transition from a substantially all composite to a substantially all metal material may reduce or eliminate the possibility of corrosion and/or problems stemming from eccentricity. During lay-up of the composite structure, relatively thin, flexible metal sheets of metal are substituted for a number of composite plies, and the transition from composite plies to metal sheets occurs at staggered locations so as to provide adequate load transfer from the composite portion to the metal portion. The staggered transition results in an interleaving between the composite plies and the metal sheets and creates multiple bond lines that may reduce the occurrence and/or propagation of cracks or disbonds in the joint. An adhesive placed between the metal sheets binds and unitizes the sheets into a nearly solid metal fitting. 
         [0009]    The composite-to-metal joint may be configured as a finger type, step lap joint in order to reduce residual stresses that may be induced in the joint during cooling of the hybrid composite structure following a thermal cure cycle. The bonded metal sheets employed in the joint form a metal laminate that may be used in a variety of other applications, and which exhibits improved performance compared to monolithic metal structures. In some applications, the composite-to-metal joint utilizing the metal laminate may be used to reinforce an edge of a composite structure or to reinforce an area of a composite structure around fasteners. Additional advantages of the disclosed composite-to metal joint may include improved joint robustness, reduced weight, improved safety, less maintenance, weight savings, improved inspectability, strength improvements, and reduced manufacturing costs. The disclosed metal laminate used in the composite-to-metal joint may enable a structure to have weight and fatigue characteristics of composite resin laminates while providing the strength and durability of a metal structure. The composite-to-metal joint may reduce or avoid the need for machined end-fittings for some composite resin structure applications. A shorter bond length resulting from use of the disclosed joint may minimizes residual (or cured in) stresses due to CTE (coefficient of thermal expansion) mismatch between the metallic and composite materials forming the joint, and may also benefit the in-service performance of the joint where service temperatures can vary 225 degrees F. or more. 
         [0010]    According to one disclosed embodiment, a metal structure is provided that exhibits improved strain performance. The metal structure comprises at least a first metal laminate including a first plurality of metal sheets bonded together. The metal structure further comprises a plurality of layers of a bonding adhesive forming adhesive bonds between the metal sheets. The metal laminate includes at least one through hole therein adapted to receive a fastener. The metal structure may further comprise a second metal laminate including a second plurality of metal sheets bonded together, and at least one fastener joining the first and second metal laminates together. 
         [0011]    According to another disclosed embodiment, an integrated attachment fitting is provided for a structure. The attachment fitting comprises a composite resin portion, a metal portion, and a composite-to-metal joint between the composite resin portion and the metal portion. The composite resin portion includes a plurality of fiber reinforced resin plies, and the metal portion includes a plurality of metal sheets bonded together. The composite-to-metal joint includes overlapping steps between the fiber reinforced resin plies and the metal sheets. The composite-to-metal joint may comprise a finger joint. In one application, the structure may comprise an aircraft vertical stabilizer, and the metal portion may be a metal laminate attachment lug having a through-hole therein adapted to receive a bolt for attaching the lug to an aircraft fuselage. The composite resin portion forms part of the aircraft vertical stabilizer. In another application, the structure may be an aircraft wing, and the metal portion is a metal laminate having a plurality of through-holes therein adapted to receive fasteners for attaching the wing to a center wing box on an aircraft fuselage. The composite-to-metal joint may be one of a finger lap joint, a tapered lap joint, a vertical lap joint, and a lap joint having a variable overlap. In a further application, the structure may be a rotor blade having a root adapted to be attached to a rotating hub, and the metal portion includes a metal laminate located at the root, wherein the metal laminate has a through-hole therein adapted to receive a retention bolt for retaining the rotor blade on the rotating hub. In still another application, the composite-to-metal joint is an overlapping splice joint adapted to join two fuselage sections of an aircraft. 
         [0012]    According to a further embodiment, a fastener reinforcement is provided for reinforcing an area of a multi-ply composite structure. The fastener reinforcement comprises a metal laminate including a plurality of metal sheets bonded together, wherein the metal laminate has a through-hole adapted to receive a fastener therein. The fastener reinforcement further comprises a composite-to-metal joint between the metal laminate and the composite structure. The meal sheets have edges that are interleafed with the plies of the composite structure. 
         [0013]    According to another disclosed embodiment, a method is provided of fabricating a composite structure, comprising assembling at least a first stack of metal sheets, and laminating the first stack of metal sheets together by placing a layer of adhesive between each of the metal sheets. The method further comprises assembling a second stack of metal sheets, laminating the second stack of metal sheets together by placing a layer of adhesive between each of the metal sheets, and fastening the first and second stacks of metal sheets by passing fasteners through the first and second stacks of metal sheets. 
         [0014]    According to still another embodiment, a method is provided of reinforcing an area of a composite laminate containing a fastener passing through the thickness of the composite laminate. The method comprises integrating a multi-ply metal laminate into the area of the composite laminate to be reinforced, and forming a through-hole in the metal laminate for receiving the fastener. Integrating the metal laminate is performed by interleafing plies of the metal laminate with plies of the composite laminate to form a finger joint between the metal laminate and the composite laminate. According to a further disclosed embodiment, a method is provided of reinforcing an edge of a multi-ply fiber reinforced resin laminate. The method comprises joining a metal laminate to the resin laminate along the edge of the resin laminate. Joining the metal laminate to the resin laminate is performed by interleafing edges of the plies of the metal laminate and the resin laminate. The interleafing may be performed in a manner to form a finger joint between the metal laminate and the resin laminate. 
     
    
     
       BRIEF DESCRIPTION OF THE ILLUSTRATIONS 
         [0015]      FIG. 1  is an illustration of a sectional view of a composite structure having a composite-to-metal joint. 
           [0016]      FIG. 2  is an illustration of a perspective view of the composite structure including the composite-to-metal joint. 
           [0017]      FIG. 3  is an illustration of a perspective view of the area designated as  FIG. 3  in  FIG. 2 . 
           [0018]      FIG. 4  is an illustration of a cross sectional view of the joint, better showing interleaving between composite plies and the metal sheets. 
           [0019]      FIG. 5  is an illustration of a cross sectional view of two separated layers of the joint shown in  FIG. 4 , also showing the application of a film adhesive on the metal sheets. 
           [0020]      FIG. 6  is an illustration of an enlarged, cross sectional view of a portion of the joint formed by the two layers shown in  FIG. 5 . 
           [0021]      FIG. 7  is an illustration of a broad flow diagram of a method of making a composite structure having the composite joint shown in  FIGS. 2-4 . 
           [0022]      FIG. 8  is an illustration of a flow diagram showing additional details of the method shown in  FIG. 7 . 
           [0023]      FIG. 9  is a flow diagram of another method of making a composite structure having the composite joint shown in  FIGS. 2-4 . 
           [0024]      FIG. 10  is an illustration of a perspective view of a composite-to-metal finger joint having a relatively shallow double taper. 
           [0025]      FIG. 11  is an illustration similar to  FIG. 10  but showing a composite-to-metal finger joint having a relatively steep taper. 
           [0026]      FIG. 12  is an illustration of a sectional view of a composite-to-metal joint having a single taper. 
           [0027]      FIG. 13  is an illustration similar to  FIG. 12  but illustrating a composite-to-metal joint having a reversed single taper. 
           [0028]      FIG. 14  is an illustration of a cross sectional view of a composite-to-metal finger joint having a symmetric double taper. 
           [0029]      FIG. 15  is an illustration similar to  FIG. 14  but illustrating a symmetric reversed double taper finger joint. 
           [0030]      FIG. 16  is an illustration of a cross sectional view of a vertical composite-to-metal finger joint. 
           [0031]      FIG. 17  is an illustration of a cross sectional view of a composite-to-metal finger joint having variable overlap between the plies. 
           [0032]      FIG. 18  is an illustration of a plan view of a composite structure having a laminated metal reinforcement around a fastener. 
           [0033]      FIG. 19  is an illustration of a cross sectional view taken along the line  19 - 19  in  FIG. 18 . 
           [0034]      FIG. 20  is an illustration of an exploded, perspective view of a typical aircraft employing composite-to-metal joints. 
           [0035]      FIG. 21  is an illustration of a sectional view taken along the line  21 - 21  in  FIG. 20 , showing a typical composite-to-metal joint between fuselage sections. 
           [0036]      FIG. 22  is an illustration of a perspective view of a composite-to-metal joint between an aircraft wing and a center wing box. 
           [0037]      FIG. 23  is an illustration of a perspective view of a portion of a skin of the wing box shown in  FIG. 22 . 
           [0038]      FIG. 24  is an illustration of the area designated as  FIG. 24  in  FIG. 22 . 
           [0039]      FIG. 25  is an illustration of a perspective view of an aircraft vertical stabilizer, parts being broken away in section for clarity. 
           [0040]      FIG. 26  is an illustration of a side view showing attachment of the stabilizer shown in  FIG. 25  to a fuselage using a lug containing a composite-to-metal joint. 
           [0041]      FIG. 27  is an illustration of a side view of a forward portion of an aircraft, illustrating a hatchway reinforced by a composite-to-metal joint. 
           [0042]      FIG. 28  is an illustration of a sectional view taken along the line  28 - 28  in  FIG. 27 . 
           [0043]      FIG. 29  is an illustration of a perspective view of a helicopter. 
           [0044]      FIG. 30  is an illustration of a perspective view of a rotor assembly of the aircraft shown in  FIG. 29 . 
           [0045]      FIG. 31  is an illustration of the area designated as  FIG. 31  in  FIG. 30 . 
           [0046]      FIG. 32  is an illustration of a cross sectional view of a bonded metal laminate. 
           [0047]      FIG. 33  is an illustration of a cross sectional view of two bonded metal laminates joined together by fasteners. 
           [0048]      FIG. 34  is a flow diagram showing a method fabricating the bonded metal laminate shown in  FIG. 32 . 
           [0049]      FIG. 35  is an illustration of a flow diagram of a method of reinforcing a composite laminate containing a fastener. 
           [0050]      FIG. 36  is an illustration of a flow diagram of aircraft production and service methodology. 
           [0051]      FIG. 37  is an illustration of a block diagram of an aircraft. 
       
    
    
     DETAILED DESCRIPTION 
       [0052]    Referring first to  FIG. 1 , a hybrid composite structure  20  includes a composite resin portion  22  joined to a metal portion  24  by a transition section  25  that includes a composite-to-metal joint  26 . In the illustrated example, the composite structure  20  is a substantially flat composite sheet, however depending upon the application, the structure  20  may have one or more curves, contours or other geometric features. For example, composite structure  20  may comprise an inner and/or outer contoured skin  20  of an aircraft (not shown) which is secured to a frame portion  28  of the aircraft by means of a lap joint  30  and fasteners  32  which pass through the composite structure  20  into the frame portion  28 . 
         [0053]    The frame portion  28  may comprise a composite, a metal or other rigid material, and the metal portion  24  of the structure  20  may serve as a rigid metal fitting  24  that is suited to transfer a range of loads and types of loadings between the frame portion  28  and the composite portion  20 . As will be discussed below in more detail, the metal portion  24  may comprise any of various metals such as, without limitation, titanium that is substantially non-reactive to and compatible with the composite portion  22  and the frame portion  28 . In one practical embodiment for example, and without limitation, the composite resin portion  22  may comprise a carbon fiber reinforced epoxy, the metal portion  24  may comprise a titanium alloy, and the frame  28  may comprise an aluminum alloy or a composite. The transition section  25  and the joint  26  are strong enough to carry the typical range and types of loads between the composite resin portion  22  and the metal portion  24 , including but not limited to tension, bending, torsion and shear loads. Although the illustrated transition section  25  and joint  26  are formed between an all composite resin portion  22  and the all metal portion  24 , it may be possible to employ them to join two differing composite structures (not shown) or two differing metal structures (not shown). 
         [0054]    Referring to  FIGS. 1-4 , a layup of composite material plies  35  is terminated at a interface location  39  referred to later herein as a transition point  39 , where a metal sheet or ply  37  of the substantially the same thickness as the composite material plies  35  continues to the metal edge  24   a  of the metal portion  24 , and the layup is repeated with a composite-to-metal interface  39  that is staggered toward the metal edge  24   a  from the prior interface location  39  and includes a ply of structural metal adhesive  45  (see  FIGS. 5 and 6 ) between the metal plies  37 , with the next composite-to-metal interface  39  staggered away from the metal edge  24   a  to produce a nested splice  27 . This staggered interface stacking, which produces nested tabs  29  (see  FIG. 3 ), is continued to the full thickness of the hybrid composite structure  20  with none of the composite plies  35  extending fully to the metal edge  24   a  of the all metal portion  24   
         [0055]    Referring now also to  FIGS. 2-4 , the composite portion  22  of the structure  20  comprises a laminated stack  34  of fiber reinforced resin plies  35 , and the metal portion  24  of the structure  20  comprises a stack  36  of metal sheets or plies  37  that are bonded together to form a laminated, substantially unitized metal structure. As shown in  FIGS. 5 and 6 , the composite plies  35  and the metal sheets  37  are arranged in layers  38 . Each of the layers  38  comprises one or more of the composite plies  35  in substantially edge-to-edge abutment with one of the metal sheets  37 . Thus, each of the layers  38  transitions at a point  39  from a composite i.e. composite resin plies  35 , to a metal, i.e. metal sheet  37 . 
         [0056]    The transition points  39  are staggered relative to each other according to a predetermined lay-up schedule such that the plies  35  and the metal sheets  37  overlap each other in the transition section  25  ( FIG. 1 ). Staggering of the transition points  39  creates multiple bond lines that may reduce the occurrence and/or propagation of cracks or disbonds in the joint  26 . The staggering of the transition points  39  also results in a form of interleaving of the composite plies  35  and the metal sheets  37  within the joint  26  which forms a nested splice  27  between the all composite portion  22  and the all metal portion  24 . This nested splice  27  may also be referred to as a finger bond  26 , a finger joint  26  or a multiple step lap joint  26 . The adjacent ones of the transition points  39  are spaced from each other in the in-plane direction of the structure  20  so as to achieve a bonded joint  26  that exhibits optimum performance characteristics, including strength and resistance to disbonds and propagation of inconsistencies such as cracks. In the illustrated example, the nested splice  27  forming the joint  26  is a form of a double finger joint in which the transition points  39  are staggered in opposite directions from a generally central point  55  of maximum overlap. However, as will be discussed blow in more detail, other joint configurations are possible including but not limited to a single finger joint in which the multiple transition points  39  are staggered in a single direction. 
         [0057]    The composite plies  35  may comprise a fiber reinforced resin, such as without limitation, carbon fiber epoxy, which may be in the form of unidirectional prepreg tape or fabric. Other fiber reinforcements are possible, including glass fibers, and the use of non-prepreg materials may be possible. The composite plies  35  may have predetermined fiber orientations and are laid up according to a predefined ply schedule to meet desired performance specifications. As previously mentioned, the bonded sheets  37  may comprise a metal such as titanium that is suitable for the intended application. In the illustrated example, the stack  36  of metal sheets  37  has a total thickness t 1  which is generally substantially equal to the thickness t 2  of the laminated stack  34  of plies  35 . In the illustrated example however, t 2  is slightly greater than t 1  by a factor of the thickness of several overwrap plies  43  on opposite sides of the stack  37 . 
         [0058]    The use of a multiple step lap joint  26  may increase the bond area along the length of the transition section  25 , compared to a scarf type joint or other types of joints which may require a longer length transition section  25  in order to achieve a comparable bond area between the composite resin portion  22  and the metal portion  24 . Following thermal curing, cooling of the hybrid composite structure  20  may result in residual stresses in the joint  26  due to a mismatch between the coefficient of thermal expansion (CTE) of the composite resin portion  22  and the metal portion  24 . The amount of thermal expansion during curing is a function of the CTE of the composite resin portion  22  and the metal portion  24 , as well as the length of the transition section  25 . Use of the step lap joint  26 , rather than a scarf type or other type of joint may reduce the amount of these residual stresses because of the reduction in the length of the transition section  25  that is needed to obtain a preselected amount of bond area between the two portions  22 ,  24  of the joint  26 . Reduction of the length of the transition section  25  may also reduce residual stresses in the joint  26  after the aircraft is placed in service where large temperature extremes may be encountered during either normal or extreme operations. 
         [0059]      FIGS. 5 and 6  illustrate details of two adjoining layers  38  of the joint  26  shown in  FIGS. 2-4 . In this example, each layer  38  comprises four plies  35  having a collective total thickness T 1 . The individual metal sheets  37  of the adjacent layers  38  are bonded together by means of a layer of structural adhesive  45 , which may comprise a commercial film adhesive or other forms of a suitable adhesive that is placed between the metal sheets  36  during the lay-up process. 
         [0060]    The combined thickness of each metal sheet  37  and one layer of adhesive  45  represented as T 2  in  FIG. 5  is substantially equal to the thickness T 1  of the composite plies  35  in the layer  38 . Although not shown in the Figures, a thin film of adhesive may be placed between the plies  35  to increase the interlaminar bond strength. In one practical embodiment, titanium alloy metal sheets may be used which each have a thickness of approximately 0.0025 inches, the film adhesive  45  may be approximately 0.005 inches thick, and four composite carbon fiber epoxy plies  35  may be used in each layer  38  having a collective total thickness of about 0.30 inches. Depending on the application, the use of metals other than titanium may be possible. The distance between adjacent transition points  39 , and thus the length of the overlap between the layers  38 , as well as the thickness and number of composite plies  35  and the thickness of the metal sheets  37  will depend on the requirements of the particular application, including the type and magnitude of the loads that are to be transmitted through the joint  26 , and possibly other performance specifications. It should be noted here that the bonded metal sheets  37  is not limited to use in a composite metal joint  26  discussed above. As will be discussed later below, a metal structure comprising bonded metal sheets  37  has a variety of other applications because of the superior strain performance it may exhibit, compared to monolithic metal structures. 
         [0061]    The differing layers  38  of the joint  26  between the two differing materials of the composite and metal portions  22 ,  24  respectively ( FIG. 1 ), render the structure  20  well suited to nondestructive evaluations of bond quality using embedded or mounted sensors (not shown). Ultrasonic structural waves (not shown) may be introduced into the structure  20  at the edge of the metal portion  24 , at the composite portion  22  or in the transition section  25 . These ultrasonic waves travel through what amounts to a waveguide formed by the metal sheets and the interfaces (not shown) between the composite plies  35  and the metal sheets  37 . MEMS-based (microelectromechanical) sensors, thin piezo-electric sensors (not shown) or other transducers placed in the structure  20  may be used to receive the ultrasonic structural waves for purposes on analyzing the condition of the bondlines in the joint  26 . 
         [0062]    Referring now to  FIG. 7 , one method of making the composite structure  20  comprises forming a multi-layer composite lay-up as shown at  65 . Forming the lay-up includes laying up a composite resin portion  22  at step  67 , and laying up a metal portion  24  at  69 . The step  65  of forming the layup further includes forming a composite-to-metal joint between the composite resin portion and the metal portion of the lay-up, shown at  71 . 
         [0063]      FIG. 8  illustrates additional details of the method shown in  FIG. 7 . Beginning at step  40 , individual metal sheets  37  are trimmed to a desired size and/or shape. Next at  42 , the surfaces of the metal sheets  37  are prepared by suitable processes that may include cleaning the sheets  37  with a solvent, drying them, etc. Then at  44 , the lay-up is assembled by laying up the metal sheets  36  and the composite plies  35  in a sequence that is determined by a predefined ply schedule (not shown) which includes a predetermined staggering of the transition points  39  between the plies  35  and the metal sheet  37  in each layer  38 . 
         [0064]    During the lay-up process, the metal sheets  37  are sequenced like plies into the lay-up, much like composite plies are sequenced into a lay-up in a conventional lay-up process. As shown at step  46 , adhesive may be introduced between the metal sheets  37  in order to bond them together into a unitized metal structure. Similarly, although not shown in  FIG. 8 , a bonding adhesive may be introduced between the individual composite plies  35  in order to increase the bond strength between these plies  35 . Next, at  48 , the lay-up may be compacted using any of several known compaction techniques, such as vacuum bagging following which the lay-up is cured at step  50  using autoclave or out-of-autoclave curing processes. At step  52 , the cured composite structure  20  may be trimmed and/or inspected, as necessary. 
         [0065]      FIG. 9  illustrates still another embodiment of a method of making a hybrid composite part  20 . The method begins at step  73  with laying at least one composite ply  35  that is terminated at an interface location  39  on a suitable layup tool (not shown). At  75 , an adjacent metal ply  37  is laid up which is substantially the same thickness as the adjacent composite material ply  35 . As shown at  77 , the layup process is repeated with a composite-to-metal interface  39  that is staggered toward the metal edge  24   a  of the part  20  from the transition point  39 . A  79 , a ply  45  of structural adhesive is laid between the metal plies  37 . Steps  73 - 79  are repeated successively to produce a nested splice  27  and a staggered interface stacking forming nested tabs  29  to the full thickness of the hybrid part  20 , with none of the composite plies  35  extending fully to the metal edge  24   a  of the part  20 . Although not shown in  FIG. 9 , the completed layup is vacuum bagged processed to remove voids, and is subsequently cured using any suitable curing method. 
         [0066]    The composite-to-metal joint  26  previously described may be constructed in any of a variety of joint configurations in which the composite material plies  35  are interleafed with the metal plies  37 . For example, referring to  FIG. 10 , the transition section  25  of the hybrid composite structure  20  may include a composite-to-metal joint  26  having a relatively shallow taper resulting from lengths L of overlap between the composite and metal plies  35 ,  37  that are relatively long. In the example shown in  FIG. 10 , the composite-to-metal joint  26  is a double tapered finger joint. In comparison, as shown in  FIG. 11 , shorter lengths L of the overlap between the composite and metal plies  35 ,  37  results in a double tapered finger joint  26  that has a relatively steep taper, in turn resulting in a shorter transition section  25  between the composite resin and metal portions  22 ,  24  respectively. The length L of the overlap may be optimized for the particular application. 
         [0067]      FIGS. 12-17  illustrate other examples of composite-to-metal joint  26  configurations. In one alternative, the composite-to-metal joint  26  may comprise a double tapered finger joint  26  that includes a tapered or layered multi-ply construction above and below a composite-to-metal interface  39 , wherein one or more overlap lengths, e.g., lengths L, may be chosen or optimized relative to a particular real estate constraint, area, or transitional stress or strain requirement. In one example, the real estate constraint or area may require a shorter transition section, for instance, between the composite resin and metal portions. In some applications, a transitional stress or strain requirement may require progressively less stress or strain along a portion of the structure. For example,  FIG. 12  illustrates a single taper lap joint  26 , while  FIG. 13  illustrates a single reverse taper lap joint  26 . In  FIG. 14 , the joint  26  is configured as a double tapered, substantially symmetrical, staggered finger lap joint while  FIG. 15  illustrates a reverse double tapered finger lap joint  26 . The use of the staggered finger lap joints  26  shown in  FIGS. 14 and 15  may be preferred in some applications because the joint may have a CTE interface that is less than an equivalent step lap joint of a longer transition section  25  ( FIG. 10 ). In  FIG. 16 , the composite-to-metal joint  26  takes the form of a vertical lap finger joint, while  FIG. 17  illustrates a composite-to-metal joint  26  in which the overlap between the composite and the metal plies  35 ,  37  is variable through the thickness of the joint  26 . 
         [0068]    Attention is now directed to  FIGS. 18 and 19  which illustrate a hybrid composite structure  20  comprising a composite resin portion  22  and a metal portion  24  that forms a metal laminate reinforcement  76  around a fastener passing through the hybrid composite structure  20 . The metal portion  24  forming the metal laminate reinforcement  76  comprises a stack  36  of metal sheets or plies  37  that are bonded together, similar to the metal laminates previously described. The metal laminate reinforcement  76  is connected to the surrounding composite resin portion  22  by a circumferential composite-to-metal joint  26 , as shown in  FIG. 19  which, in the illustrated embodiment, comprises a double tapered finger lap joint, similar to that shown in  FIGS. 4 ,  10 , and  14 . In one alternative, staggered finger lap joints may include a transition region where one or more edges of composite material plies, metal plies, or combinations thereof may have varying levels of overlap or non-overlap to achieve or meet a desired CTE interface coefficient, a desired real estate constraint, an area constraint, or transitional stress or strain requirement. In one example, real estate constraint or area may require a shorter transition section, for instance, between the composite resin and metal portions or metal plies. In one example, transitional stress or strain requirement may require progressively less stress or strain along a portion of the structure. 
         [0069]    The metal laminate reinforcement  76  includes a central through-hole  85  through which the fastener  78  passes. The fastener  78  may comprise for example and without limitation, a bolt or rivet  78  having a body  78   a  and heads  78   b  and  78   c.  Although not shown in the drawings, the fastener  78  may be used to attach a structure to the composite structure  20 , or to secure the hybrid composite structure  20  to another structure. The metal laminate reinforcement  76  functions to strengthen the area surrounding the fastener  78  and may better enable the composite structure  20  to carry loads in the area of the fastener  78 . 
         [0070]    The composite-to-metal joint  26  previously described may be employed in a variety of applications, including those in the aerospace industry to join composite structures, especially in areas where a composite structure is highly loaded. For example, referring to  FIG. 20 , an airplane  80  broadly comprises a fuselage  82 , left and right wings  84 , a vertical stabilizer  92  and a pair of horizontal stabilizers  94 , and a wing box  108 . The airplane  80  may further include a pair of engines  88  surrounded by engine nacelles  86 , and landing gear  90 . 
         [0071]    The composite-to-metal joint  26  previously described may be employed to join or mount any of the components shown in  FIG. 20 . For example, composite-to-metal joints  26  may be employed to mount the wings  84  on the center wing box  108 , as will be discussed below in more detail. Similarly, a composite-to-metal joint  26  may be employed to attach the vertical stabilizer  92  and/or the horizontal stabilizers  94  to the fuselage  82 . The composite-to-metal joints  26  may be employed to mount the landing gear  90  on the wings  84 , as well as to mount engines  88  and engine nacelles  86  on pylons (not shown) on the wings  84 . Further, the disclosed composite-to-metal joint  26  may be employed to join fuselage sections  82   a  together. For example, referring to  FIGS. 20 and 21 , fuselage sections  82   a  may be joined together by a co-bonded lap joint indicated at  96 , wherein each of the adjoining fuselage sections  82   a  comprises a metal laminate stack  36  and finger overlaps  98 ,  100  between composite resin and metal plies  35 ,  37  respectively. In this example, the metal laminate stacks  36  of the respective fuselage sections  82   a  may be joined together, as by bonding using a suitable bonding adhesive. 
         [0072]    Referring now to  FIG. 22 , each of the wings  84  ( FIG. 20 ) may be attached to the center wing box  108  by an attachment joint, generally indicated at  104 . Each of the wing  106  and the wing box  108  broadly comprises an outer skin  120  attached to spanwise extending spars  110 . The attachment joint  104  includes an attachment fitting  114  having a pair of flanges  118  that are attached by bolts  122  or other suitable fasteners to the skins  120 . The attachment joint  104  may be reinforced by C-shaped channels  112  and brackets  116 . 
         [0073]    Referring also now to  FIGS. 23 and 24 , each of the skins  120  includes a metal portion  24  that also forms an integrated attachment fitting which is connected to a composite resin portion  22  by a composite-to-metal joint  26  of the type previously described. Although not shown in  FIGS. 23 and 24 , the metal portion  24  of the joint  26  is formed by laminated metal plies  37 , and the composite resin portion of the joint  26  is formed by laminated composite resin plies  35 . As particularly shown in  FIG. 24 , the metal portion  24  of the joint  26  may be scarfed at  128  to receive one of the flanges  118  therein. Metal portions  24  include through-holes  124  that are aligned with the through-holes  126  in the flanges  118  of the fitting  116 . It may thus be appreciated that attachment joint  104  is reinforced by the presence of the metal portions  24  which are attached to the metal attachment fitting  114  by the bolts  122 . 
         [0074]      FIGS. 25 and 26  illustrate another application of composite-to-metal joint  26  that may be employed to attach a vertical stabilizer  92  or similar airfoil to an aircraft fuselage  82 . As shown in  FIG. 25 , the vertical stabilizer  92  may comprise a series of generally upwardly extending spars  130  connected with ribs  132 . A series of attachment lugs  134  on the bottom of the stabilizer  92  are each attached to mounting ears  138  on the fuselage  82  by means of attachment bolts  136  received within bushings  140  in the lugs  134 . Each of the lugs  134  comprises a fiber reinforced composite resin portion  22  and a metal portion  24  which may comprise a metal laminate. The composite resin portion  22  is joined to the metal portion by a composite-to-metal joint  26  of the type previously described. It may thus be appreciated that while the lug  134  is lightweight because of its predominantly composite construction, the area at which the lug  134  is attached to the fuselage  82  comprises a metal portion  24  which has a load bearing capacity that may be greater than the composite resin portion  22 . 
         [0075]    Attention is now directed to  FIGS. 27 and 28  which illustrate the use of a composite-to-metal joint  26  employed to reinforce the edges  142  of a fiber reinforced composite resin structure, which in the illustrated example comprises the skin  120  of an aircraft  80 . In this example, a fuselage hatch  141  has a periphery  142  terminating in an edge  144  ( FIG. 28 ) that is reinforced by a metal portion  24  comprising a metal laminate stack  36 . The metal portion  24  is joined to the composite skin  120  by a composite-to-metal joint  26 , of the type previously described. In this example, the edge  24   a  of the metal portion  24  defines the fuselage hatch  141  opening. The composite-to-fiber joint  26  may also be used to reinforce the skin  120  around other openings, such as cockpit windows  125  and passenger windows  127 . 
         [0076]    Referring now to  FIG. 29 , the composite-to-metal joint  26  may be employed to attach components on other types of aircraft, such as, for example and without limitation, a helicopter  146 . The helicopter  146  includes a main rotor assembly  148  and a tail rotor assembly  150 . The main rotor assembly  148  includes a plurality of main rotator blades  152 , and the tail rotor assembly  150  comprises a plurality of tail rotor blades  154 . Each of the main rotor blades  152  is mounted on a rotor hub  156  secured to a rotating mast  168  that is powered by one or more engines  160 . Referring particularly to  FIGS. 30 and 31 , each of the main rotor blades  152  is attached to the hub  156  by means of blade grips  164 . The root  162  of each blade  152  is held on the blade grips  164  by retention bolts  166 . Each of the blades  152  includes an elongate outer composite resin portion  22  which may be a carbon fiber epoxy composite, and a metal portion  24  that is attached to the blade grips  164  by the retention bolts  166 . Metal portion  24  of the blade  152  is connected to the outer composite resin portion by a composite-to-metal joint  26  of the type previously described. The tail rotor blades  154  shown in  FIG. 29  may similarly be attached to the tail rotor assembly  150  by a composite-to-metal joint  26 . 
         [0077]    Referring to  FIG. 32 , a metal laminate  170  comprises a plurality of generally flexible metal sheets or plies  37  which are bonded together by layers  45  of a suitable adhesive to form a structure that may exhibit performance properties that are superior to a comparable monolithic metal structure. The layers  45  of adhesive may comprise a conventional film-type structural adhesive. The metal plies  37  may be formed of the same metal or may be formed of differing metals, depending on the particular application. When the metal laminate  170  is placed in tension  175 , the tension load is individually directed to each of the metal laminate plies  37 , thereby distributing the tension load generally evenly throughout the metal structure  170 . Thus, in the event of an irregularity or inconsistency in one of the metal plies  37  that may reduce the load carrying ability of the ply  37 , the reduction is limited to that particular ply and the applied tension load is redistributed to the remaining metal plies  37  which provide strain relief. In other words, sensitive areas (i.e. plies  37 ) of the metal laminate  170  that are under load locally strain and transfer the load to adjacent metal plies  37 , resulting in a form of a progressive loading of the metal laminate  170 . 
         [0078]    The metal laminate  170  shown in  FIG. 32  may be employed to form composite-to-metal joints  26  of the type previously described, but may have other applications as well. For example, referring to  FIG. 33 , two generally flat metal laminates  170   a,    170   b  may be attached to each other by a lap joint  172  and fasteners  178  that pass through through-holes  173  the metal laminates  170   a,    170   b . The lap joint  172  employing may exhibit characteristics that are superior to joints employing monolithic structures. The metal laminates  170   a,    170   b  may form the edges of a composite structure to which the metal laminates  170   a,    170   b  are joined by composite-to-metal joints  26  of the type previously described. 
         [0079]    Referring to  FIG. 34 , a method of fabricating a structure begins at  180 , with assembling at least a first stack  36  of metal sheets or plies  37 . The metal sheets or plies  37  are then laminated together at  182  by placing a layer of structural adhesive between the sheets or plies  37  which bonds and laminates the sheets or plies  37  together into a first metal laminate  170   a.  Then, optionally at  184 , a second stack of metal sheets or plies  37  is assembled and laminated together at  186  into a second metal laminate  170   b.  At  188 , one or more through-holes  173  are formed in the first and second laminates  170   a,    170   b.  At  190 , fasteners are installed in the though-holes  173  to fasten the metal laminates  170   a ,  170   b  together. 
         [0080]    Referring to  FIG. 35 , selected areas of a fiber reinforced composite resin laminate structure may be reinforced by a method that begins at step  192  with assembling a metal laminate reinforcement  76 . At step  194 , composite resin plies  35  of the composite resin laminate structure are interleafed with the metal laminate plies  37  of the metal laminate reinforcement  76  to form a composite-to-metal step lap joint  26  in the area of the composite resin laminate structure to be reinforced. As previously discussed, the metal laminate reinforcement  76  may be used to reinforce an edge of the composite resin laminate structure, or to provide a metal reinforced area around a fastener  78 . Thus, optionally, at step  196 , a through-hole  85  may be formed in the metal reinforcement  76 , and at  198 , a fastener  78  may be installed in the through-hole  85 . 
         [0081]    Embodiments of the disclosure may find use in a variety of potential applications, particularly in the transportation industry, including for example, aerospace, marine and automotive applications. Thus, referring now to  FIGS. 36 and 37 , embodiments of the disclosure may be used in the context of an air10raft manufacturing and service method  200  as shown in  FIG. 36  and an aircraft  202  as shown in  FIG. 37 . Aircraft applications of the disclosed embodiments may include, for example, a wide variety of structural composite parts and components, especially those requiring local reinforcement and/or the use of fasteners during the assembly process. During pre-production, exemplary method  200  may include specification and design  204  of the aircraft  202  and material procurement  206 . During production, component and subassembly manufacturing  208  and system integration  210  of the aircraft  202  takes place. Thereafter, the aircraft  202  may go through certification and delivery  212  in order to be placed in service  214 . While in service by a customer, the aircraft  202  is scheduled for routine maintenance and service  216 . 
         [0082]    Each of the processes of method  200  may be performed or carried out by a system integrator, a third party, and/or an operator (e.g., a customer). For the purposes of this description, a system integrator may include without limitation any number of aircraft manufacturers and major-system subcontractors; a third party may include without limitation any number of vendors, subcontractors, and suppliers; and an operator may be an airline, leasing company, military entity, service organization, and so on. 
         [0083]    As shown in  FIG. 37 , the aircraft  202  produced by exemplary method  200  may include an airframe  218  with a plurality of systems  220  and an interior  222 . Examples of high-level systems  220  include one or more of a propulsion system  224 , an electrical system  226 , a hydraulic system  228 , and an environmental system  230 . Any number of other systems may be included. The disclosed method may be employed to fabricate parts, structures and components used in the airframe  218  or in the interior  222 . Although an aerospace example is shown, the principles of the disclosure may be applied to other industries, such as the marine and automotive industries. 
         [0084]    Systems and methods embodied herein may be employed during any one or more of the stages of the production and service method  200 . For example, parts, structures and components corresponding to production process  208  may be fabricated or manufactured in a manner similar to parts, structures and components produced while the aircraft  200  is in service. Also, one or more apparatus embodiments, method embodiments, or a combination thereof may be utilized during the production stages  208  and  210 , for example, by substantially expediting assembly of or reducing the cost of an aircraft  200 . Similarly, one or more of apparatus embodiments, method embodiments, or a combination thereof may be utilized while the aircraft  202  is in service, for example and without limitation, to maintenance and service  216 . 
         [0085]    Although the embodiments of this disclosure have been described with respect to certain exemplary embodiments, it is to be understood that the specific embodiments are for purposes of illustration and not limitation, as other variations will occur to those of skill in the art.