Abstract:
Technologies for reducing the risk of disbonding between two bonded members at areas of differing strain in the members are provided. Where the difference in strain between the two bonded members becomes excessing, movement of the members relative to each other may cause a traditional rigid adhesive to fail, resulting in a disbond or delamination in one or both composite members at the point of differing strain. Disbonding may be minimized or prevented by placing a compliant material, such as a thin sheet of rubber, between the structural members at the point of differing strain. The compliance of the material may allow the compliant material to deform enough to remain bonded to both structural members when subjected to a stress that might otherwise disbond or delaminate the bonded members.

Description:
BACKGROUND 
     Two semi-rigid parts that are bonded together may have a risk of disbonding if the two parts each experience different levels of strain at a same point along the bond between the parts. For example, a skin of a wing structure of an aircraft may have a number of structural members, or “stringers,” joined to one surface in order to provide shape, strength, and rigidity to the skin. The strains experienced by the skin and the stringer may be the same at most points along the wing structure. However, at a point where the stringer terminates, the skin and the stringer may experience differing levels of strain when the wing structure is loaded, such as during flight. If the stringer is joined to the skin in a traditional method, such using rivets or bolts, the differing strains in the stringer and skin may be tolerated by the joint, since the members are allowed to “fret” or move relative to each other without the joint failing. 
     If the stringer is bonded to the skin using a rigid material, however, such as an adhesive that is hard or brittle when it cures, then the difference in strains between these two members may cause the adhesive to fail and the stringer to disbond from the skin. Alternatively, the adhesive may hold, but the differing strains may cause a delamination of the layers of a composite skin and/or stringer. Such a disbond or delamination may cause fuel leakage from the wing tanks, excessive aircraft noise during flight, weakening of the wing structure, and the like. Different solutions have been implemented to reduce the risk of disbonding between the stringer and the skin, such as softening the stringer, adding additional bolts or fasteners to the joint, tapering or feathering the stringer at the point of differing strains, and the like. However, none of these solutions sufficiently reduces the risk of disbond between the members when the wing structure is under load. 
     It is with respect to these and other considerations that the disclosure made herein is presented. 
     SUMMARY 
     It should be appreciated that this Summary is provided to introduce a selection of concepts in a simplified form that are further described below in the Detailed Description. This Summary is not intended to be used to limit the scope of the claimed subject matter. 
     Methods, structures, and systems are described herein for reducing the risk of disbonding between two bonded structural members at areas of differing strains in the members. Where the difference in strains between the two bonded members becomes excessive, movement of the members relative to each other may cause a traditional rigid adhesive to fail, resulting in a disbond and/or a delamination in the layers of composite members at the point of differing strain. Disbonding may be minimized or prevented by placing a compliant material, such as a thin sheet of rubber, between the structural members at the point of differing strain. The compliance of the material may allow the compliant material to deform enough to remain bonded to both structural members when subjected to a stress that might otherwise disbond or delaminate the bonded members. 
     According to one aspect, a structure comprises two structural members bonded together with a piece of compliant material disposed between the two structural members at an area of the bond between the members where a differential in strain may occur between the two structural members when the structure is under a load. In another aspect, a method for reducing the risk of disbond between two bonded structural members comprises identifying an area of the bond between the structural members where a differential in strain may occur between the members when placed under a load and disposing a piece of compliant material between the two structural members at the identified area of the bond. In a further aspect, a system for reducing the risk of disbond between a skin of a wing structure in an aircraft and a stringer bonded to the skin includes a piece of compliant material disposed between the skin and the stringer at an area of the bond between them where a differential in strain occurs when the wing structure is under load. 
     The features, functions, and advantages discussed herein can be achieved independently in various embodiments of the present disclosure or may be combined in yet other embodiments, further details of which can be seen with reference to the following description and drawings. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1A  is a front view of an aircraft having a front cross section area designated, and  FIG. 1B  is a cutaway diagram of a side-of-body joint of a wing structure according to of  FIG. 1A , according to embodiments presented herein. 
         FIG. 2  is a side view of a stringer attached to the skin of the wing structure, according to embodiments presented herein. 
         FIG. 3  is another side view of a stringer attached to the skin of a wing structure, according to further embodiments presented herein. 
         FIG. 4  is a flow diagram illustrating one method for reducing the risk of disbonding between two bonded members at areas of differing strain, according to the embodiments described herein. 
     
    
    
     DETAILED DESCRIPTION 
     The following detailed description is directed to technologies for reducing the risk of disbonding between two bonded members at areas of differing strain in the members. As utilized herein, “disbonding” refers to the breaking of the adhesive or bond between two-co-bonded members, also known as “mode 1, mode 2, or mode 3 cracking,” as well as delamination of the layers of one or both of the composite members. Further, while the embodiments of the disclosure are described herein in the context of a stringer bonded to a skin of a wing or other structure in an aircraft, it will be appreciated that embodiments of the disclosure are not limited to such applications, and that the techniques described herein may also be utilized to prevent disbonding between bonded structural members in other applications. For example, embodiments may be applicable to bonds between structural members in the root of a composite helicopter blade, the centerline joint of a horizontal tail, engine mounts, landing gear assemblies, and the like. The embodiments described herein may also be applicable in any other aircraft and non-aircraft applications comprising joints between parts where there is a disparity between the strains on the separate parts of the joint. 
     In the following detailed description, references are made to the accompanying drawings that form a part hereof, and that show, by way of illustration, specific embodiments or examples. The drawings herein are not drawn to scale and the relative proportions of the various elements may be exaggerated to illustrate aspects of the disclosure. Like numerals represent like elements throughout the several figures. 
       FIG. 1A  and  FIG. 1B  show an example configuration between two structural members where a risk of disbonding between two members may occur. Specifically,  FIG. 1A  and  FIG. 1B  show a cutaway view of a side-of-body joint  100  of a wing structure  102  to another structure in the fuselage  104  of an aircraft, such as a center wing box  106 , as shown in the figure. The wing structure  102  comprises a skin  108 . The skin  108  of the wing structure  102  may be made of a semi-rigid material, such as carbon-reinforced plastic or other composite or metal material. As is known in the art, one or more stringers  110  may be bonded to one surface of the skin  108  in order to provide shape, strength, and rigidity to the skin in the wing structure  102 . The stringers  110  may be made of the same composite or metal material as the skin  108 , or of a material with similar rigidity. 
     At the side-of-body joint  100 , the skin  108  and the stringer  110  may be joined to one or more structural elements  112 , such as a titanium cord that runs the length of the joint. The skin  108  may be joined by one set of fasteners  114 , such as bolts, rivets, or screws, to one portion of the titanium cord, while the stringer may be joined by another set of fasteners  116  to a different portion of the titanium cord. The titanium cord may be further joined to the center wing box  106  or other structure in the fuselage  104  in a similar fashion. It will be appreciated that other configurations of the side-of-body joint  100  may be conceived in which a risk of disbonding between the members may occur beyond that shown in  FIG. 1A  and  FIG. 1B . 
       FIG. 2  show further details of the wing structure  102 . According to embodiments, the stringer  110  is bonded to the skin  108  with an adhesive  202  which forms a strong but relatively rigid bond between the skin and stringer when cured. The point between the stringer  110  and the skin  108  where the adhesive  202  resides is also referred to herein as the “glueline.” Because the skin  108  and stringer  110  are made of a same material or materials with similar rigidity, the strains experienced in both the stringer and the skin while the wing structure is loaded may be substantially the same along the majority of the wing structure  102 , as shown at  204  in  FIG. 2 , where the strain E in the stringer  110  and the skin  108  are both shown as 0.004 in./in. This may represent the strain in these structural members when the wing structure  102  is under full load, such as when the aircraft comprising the wing structure is in flight or is being subjected to a maximum design stress in testing, for example. 
     However, at a point at or near the end of the glueline between the stringer  110  and skin  108 , a condition may be created where the strains experienced in one structural member is different from that in the other structural member at the same point. This condition may occur at the side-of-body joint  100  where the stringer  110  and skin  108  are attached to the structural elements  112 , or the condition may occur at the stringer “runout” towards the other end of the wing structure  102 . For example, as shown at  206 , the strain ε in the stringer  110  at the point where the glueline stops may go to zero while the strain in the skin  108  at that points remains at 0.004 in./in. This may create a relatively large in-plane displacement between the stringer  110  and the skin  108  at or near the point of differing strain levels, causing the rigid adhesive  202  to rupture or fail and resulting in a disbond between the stringer  110  and the skin  108  at that point. The differing strains may also cause a delamination to occur between the plies within a composite skin  108  or stringer  110 . 
     According to embodiments, the risk of disbonding between the structural members, e.g. the stringer  110  and the skin  108 , may be reduced or eliminated by placing a compliant interface component  208  between the stringer and the skin at the point of differing strains, as further shown in  FIG. 2 . The compliant interface component  208  may be bonded to both the stringer  110  and the skin  108  using the adhesive  202 , and the compliance of the material may allow it deform enough to remain bonded to both structural members through the relatively large displacement caused by the disparity between the respective strain levels. In addition, the compliance of the material in the compliant interface component  208  may prevent a delamination in a composite skin  108  or stringer  110  from occurring. 
     The optimal thickness of the compliant interface component  208  and the material used in its fabrication may be based on the type of adhesive  202  used at the glueline between the stringer  110  and the skin  108 , the expected difference between strain levels in the respective structural members, the environmental conditions in which the bond exists, the total area of the joint, the required load transfer driven by the design of the joint, the form factor of the chosen material, and the like. According to embodiments, the compliant interface component  208  may comprise a thin piece of rubber between 0.003 and 0.100 inches thick. In one embodiment, the compliant interface component  208  may be between 0.020 to 0.050 inches thick. This may provide sufficient compliance in the compliant interface component  208  to remain bonded to both the stringer  110  and the skin  108  under the maximum strain differential, such as 0.004 in./in, without being too thick to be inserted into the glueline. In another embodiment, the piece of rubber is further reinforced with fiberglass cloth to prevent expansion of the compliant interface component  208  during the bonding process. This may allow for more rubber to be used in the compliant interface component  208  without the rubber being squeezed out of the joint under clamp-up loads. 
     The type of rubber utilized for the compliant interface component  208  between the stringer  110  and the skin  108  may be fuel resistant to survive the fuel tank environment within the wing structure  102 , as well as heat resistant to survive the cure cycle during fabrication. The rubber may also need to retain its compliance properties at low temperatures, which may be as low as −20° to −65° F. in the wing structure when the aircraft is operating in external temperatures of down to −65° F. According to one embodiment, a fiberglass-reinforced nitrile rubber film approximately 0.030 thick is utilized for the compliant interface component  208 . In other embodiments, the compliant interface component  208  may be made of a fluoroelastomer, such as DUPONT™ VITON® from E. I. du Pont de Nemours and Company of Wilmington, Del. The compliant interface component  208  may also consist of a plastic material, such as nylon, a rubberized adhesive, a plasticized adhesive, and the like. 
     The compliant interface component  208  may occupy the full width of the bond between the stringer  110  and the skin  108 . In addition, the compliant interface component  208  may run the full length of the bond between the stringer  110  and the skin  108  where the potential difference between the strains of the respective structural members is high enough to cause possible failure of the adhesive  202 . For example, for a stringer  110  comprising an I-beam of approximately 6″ by 6″ in dimension bonded to a composite skin  108  of ½″ to ¾″ in thickness, the compliant interface component  208  may run 6″ to 12″ along the glueline from the end of the stringer, or the compliant interface component  208  may run the entire length of the stringer. 
     In one embodiment, the compliant interface component  208  is added between the stringer  110  and the skin  108  when the stringers are co-bonded to the skin  108  during fabrication of the wing structure  102 . The compliant interface component  208  may be bonded to the stringer  110  first, and then to the skin  108 , or it may be bonded to both at the same time during curing of the adhesive  202 . In another embodiment, the compliant interface component  208  may represent a change in the formulation of the adhesive  202  used at the glueline between the stringer  110  and the skin  108  at the point where the differing strains occur. The compliant interface component  208  may further be brushed onto the skin  108  and/or stringer  110  during assembly of the wing structure  102 , or may be molded into the stringer  110 . 
       FIG. 3  illustrates another example of the stringer  110  bonded to the skin  108  of the wing structure  102 , according to further embodiments. For example,  FIG. 3  may illustrate the stringer runout towards the tip of the wing structure  102 . In addition to the adhesive  202 , a number of fasteners  302 , such as bolts, screws, rivets, and the like, may be used to join the stringer  110  to the skin  108  near the end of the stringer, as well at various points along the glueline between the stringer  110  and the skin  108  of the wing structure  102 . In one embodiment, the compliant interface component  208  may replace the adhesive  202  in the area of differing strains between the stringer  110  and the skin  108  near the end of the stringer, being “bonded” to the skin and/or stringer through the natural adhesiveness of the material used in the compliant interface component. In additional embodiments, the compliant interface component  208  may be bonded to either the stringer  110  or the skin  108 , and the joint between the stringer and the skin may or may not include one or more fasteners  302 . 
       FIG. 4  shows a routine  400  for reducing the risk of disbonding between two bonded structural members, according to one embodiment. The routine  400  may be utilized to reduce the risk of disbonding between the stringer  110  and the composite skin  108  at the end of the stringer in the wing structure  102 , as described above in regard to  FIGS. 2 and 3 , for example. The routine  400  begins at operation  402 , where an area of differing strain between the bonded structural members while under load is identified. The area may be identified for points along the bonding between the structural members where the difference in levels of strain in the respective members may cause a failure of the adhesive  202  used to bond the members together. For example, a substantial difference between the strain in the stringer  110  and the skin  108  of the wing structure  102  may exist at or near the end of the stringer, such as at the side-of-body joint  100  or the stringer runout, for example. 
     From operation  402 , the routine  400  proceeds to operation  404 , where parameters of a compliant interface component are selected, such as the size, type, and thickness of the material to be used in its fabrication. As described above, the thickness of the compliant interface component  208  and the material used in its fabrication may be based on the type of adhesive  202  used at the glueline between the stringer  110  and the skin  108 , the expected difference between strain levels in the respective structural members, the environmental conditions in which the bond exists, the total area of the joint, the required load transfer driven by the design of the joint, the form factor of the chosen material, and the like. The compliant interface component  208  may further be reinforced with fiberglass cloth or other material to prevent expansion of the compliant interface component  208  during the bonding process. 
     The routine  400  proceeds from operation  404  to operation  406 , where the compliant interface component  208  is disposed between the structural members in the area of differing strain identified in operation  402 . For example, a 0.030 inch thick piece of fiberglass-reinforced nitrile rubber film may be inserted at the glueline between the stringer  110  and the skin  108  near the end of the stringer, as shown in  FIGS. 2 and 3 . The rubber film may be added between the stringer  110  and the skin  108  at the moment the stringers are co-bonded to the skin during fabrication of the wing structure  102 . The compliant interface component  208  may be bonded to one or both of the structural members, using adhesive or some other bonding method, and may occupy the full width of the joint between the members and run the full length of the area of differing strain identified in operation  402 . From operation  406 , the routine  400  ends. 
     Based on the foregoing, it should be appreciated that technologies for reducing the risk of disbonding between two bonded members at areas of differing strain in the members are provided herein. The subject matter described above is provided by way of illustration only and should not be construed as limiting. Various modifications and changes may be made to the subject matter described herein without following the example embodiments and applications illustrated and described, and without departing from the true spirit and scope of the present invention, which is set forth in the following claims.