Patent Document

TECHNICAL FIELD 
       [0001]    This disclosure generally relates to composite structures, especially to fiber reinforced resin laminates, and deals more particularly with a hybrid composite having a composite-to-metal joint. 
       BACKGROUND 
       [0002]    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. 
         [0003]    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. 
         [0004]    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. 
       SUMMARY 
       [0005]    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. 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, 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. 
         [0006]    According to one disclosed embodiment, a composite structure is provided, comprising a laminated stack of fiber reinforced resin plies and a stack of metal sheets. The metal sheets have edges that are interleaved with the edges of the fiber reinforced resin plies to form a composite-to-metal joint connecting the fiber reinforced resin plies with the metal sheets. 
         [0007]    According to another embodiment, a hybrid resin-metal structure is provided comprising a composite resin portion, a metal portion, and a transition section between the resin and metal portions. The resin portion includes laminated plies of fiber reinforced resin, and the metal portion includes bonded sheets of metal. The transition section includes staggered overlaps between the laminated plies and the metal sheets. 
         [0008]    According to another embodiment, a hybrid composite metal part is provided. The part comprises a layup of fiber reinforced composite material that is terminated at an interface location. At the interface location, a metal ply of the same thickness as the composite material continues to the metal edge of the part, and the layup is repeated with a composite to metal interface that is staggered toward the edge of the part from the prior interface location. A ply of structural adhesive is placed between the metal plies, with the next metal to composite interface staggered away from the part edge to produce a nested splice, and the staggered interface stacking produces nested tabs is continued to the full thickness of the part with none of the composite plies extending fully to the edge of the part. 
         [0009]    According to still another embodiment, a method is provided of fabricating a composite structure. The method comprises forming a multi-ply composite lay-up having a composite portion and a metal portion, and forming a composite-to-metal joint between the composite portion and the metal portion. The method further includes compacting and curing the layup. 
         [0010]    According to a further embodiment, a method is provided to produce a hybrid metal part. The method comprises laying at least one fiber reinforced composite ply that is terminated at a interface location, and laying an adjacent metal ply where the metal ply is of substantially the same thickness as the adjacent fiber reinforced composite ply. The steps of laying composite plies and adjacent metal plies are repeated to form a composite to metal interface that is staggered toward said an edge of the part from the prior interface location. The method further comprises laying a ply of structural adhesive between the metal plies, and repeating the composite and metal ply layup where the next metal to composite interface is staggered away from the part edge to produce a nested splice. 
     
    
     
       BRIEF DESCRIPTION OF THE ILLUSTRATIONS 
         [0011]      FIG. 1  is an illustration of a sectional view of a composite structure having a composite-to-metal joint. 
           [0012]      FIG. 2  is an illustration of a perspective view of the composite structure including the composite-to-metal joint. 
           [0013]      FIG. 3  is an illustration of a perspective view of the area designated as  FIG. 3  in  FIG. 2 . 
           [0014]      FIG. 4  is an illustration of a cross sectional view of the joint, better showing interleaving between composite plies and the metal sheets. 
           [0015]      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. 
           [0016]      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 . 
           [0017]      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 . 
           [0018]      FIG. 8  is an illustration of a flow diagram showing additional details of the method shown in  FIG. 7 . 
           [0019]      FIG. 9  is a flow diagram of another method of making a composite structure having the composite joint shown in  FIGS. 2-4 . 
           [0020]      FIG. 10  is an illustration of a flow diagram of aircraft production and service methodology. 
           [0021]      FIG. 11  is an illustration of a block diagram of an aircraft. 
       
    
    
     DETAILED DESCRIPTION 
       [0022]    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 of an aircraft (not shown) which is secured to a frame  28  portion of the aircraft by means of a lap joint  30  and fasteners  32  which pass through the skin  20  into the frame  28 . 
         [0023]    The frame  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  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  20  and the structure  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). 
         [0024]    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 metal-to-composite interface  39  staggered away from the part 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   
         [0025]    Referring now also to  FIGS. 2-4 , the composite portion  22  of the structure  20  comprises a laminated stack 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 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 . 
         [0026]    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  26  in which the transition points  39  are staggered in opposite directions from a generally central point  55  of maximum overlap. However, 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. 
         [0027]    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 . 
         [0028]      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. 
         [0029]    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  37  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. 
         [0030]    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  37  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 . 
         [0031]    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 . 
         [0032]      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  36  in each layer  38 . 
         [0033]    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. 
         [0034]      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 play  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 prior interface location  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 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. 
         [0035]    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. 10 and 11 , embodiments of the disclosure may be used in the context of an aircraft manufacturing and service method  60  as shown in  FIG. 10  and an aircraft  62  as shown in  FIG. 11 . Aircraft applications of the disclosed embodiments may include, for example, a wide variety of structural composite parts and components, especially those requiring the use of fasteners during the assembly process. During pre-production, exemplary method  60  may include specification and design  64  of the aircraft  62  and material procurement  66 . During production, component and subassembly manufacturing  68  and system integration  70  of the aircraft  62  takes place. Thereafter, the aircraft  62  may go through certification and delivery  72  in order to be placed in service  74 . While in service by a customer, the aircraft  62  is scheduled for routine maintenance and service  76  (which may also include modification, reconfiguration, refurbishment, and so on). 
         [0036]    Each of the processes of method  60  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. 
         [0037]    As shown in  FIG. 11 , the aircraft  62  produced by exemplary method  60  may include an airframe  78  with a plurality of systems  80  and an interior  82 . Examples of high-level systems  82  include one or more of a propulsion system  84 , an electrical system  86 , a hydraulic system  88 , and an environmental system  90 . Any number of other systems may be included. The disclosed method may be employed to fabricate parts, structures and components used in the airframe  78  or in the interior  82 . Although an aerospace example is shown, the principles of the disclosure may be applied to other industries, such as the marine and automotive industries. 
         [0038]    Systems and methods embodied herein may be employed during any one or more of the stages of the production and service method  60 . For example, parts, structures and components corresponding to production process  68  may be fabricated or manufactured in a manner similar to parts, structures and components produced while the aircraft  62  is in service. Also, one or more apparatus embodiments, method embodiments, or a combination thereof may be utilized during the production stages  68  and  70 , for example, by substantially expediting assembly of or reducing the cost of an aircraft  62 . Similarly, one or more of apparatus embodiments, method embodiments, or a combination thereof may be utilized while the aircraft  62  is in service, for example and without limitation, to maintenance and service  76 . 
         [0039]    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.

Technology Category: b