Patent Publication Number: US-2019177007-A1

Title: Composite repair kit

Description:
BACKGROUND 
     1. Field 
     The present disclosure is related to methods and systems for repairing structures that include composite materials, and in particular, to methods and systems for effecting such repairs on aircraft with limited resources and time. 
     2. Related Art 
     The use of structures constructed of composite materials have increased significantly in many areas including the aircraft industry. The reason for this includes the benefits of increased strength and rigidity, reduced weight, and the reduced number of parts per structure. However, with the increased use of structures constructed of composite materials (i.e., composite structures) comes the need to properly repair any damage to these types of structures. Specifically, large area composite repair is rapidly becoming an important support issue for aircraft that utilize composite structures. As an example, while small-sized damage to an aircraft fuselage may require scarfed repair, as damage size increases, other approaches are needed to the point that for large repairs, integration of the repair structure and the surrounding structure will require significant time to repair. Moreover, composite structures often require extensive repair work that may ground an aircraft for a significant amount of time that may be, for example, three or more weeks, thereby adding significantly to the support costs of the aircraft since the aircraft is taken out of operation. 
     Generally, the current procedure for repairing large area damage on an aircraft utilizing composite materials is described in  FIG. 1 . In  FIG. 1 , a flowchart of a known method  100  is shown. The method  100  starts  102  when a large area of damage occurs on an aircraft. The known method  100  then includes non-destructive inspection (“NDI”) of the damage to determine the size and location of the damage  104 . The information is then typically passed to a commercial aviation services department of the manufacturer of the aircraft that determines what type of repair needs to be performed and then creates and sends a request (i.e., a repair definition)  106  to the support engineering department of the manufacturer to analyze the damage and design a custom repair kit  108 . The support engineering department then designs a custom repair kit  110  and sends a request to the manufacturing department of the manufacturer to fabricate the custom repair kit. The manufacturing department then needs to dedicate tooling, materials, processes, workforce, and manufacturing facility space  112  to produce  114  the custom repair kit that may take, as an example, three or more weeks to complete. Once the custom repair kit is complete, it is shipped  116  to the location of the affected aircraft where aircraft-on-ground (“AOG”) personal deliver and complete the repair on the aircraft  118 . The method then ends  120 . 
     This known method is a time consuming and expensive process. As such, there is a need for a system and method that allows for repairing large area damage on aircraft with composite structures that is faster, more efficient, and less costly than the present approaches. 
     SUMMARY 
     Disclosed is a method for repairing a damaged portion of a composite fuselage or wing on an aircraft. The method includes performing a non-destructive inspection (“NDI”) of the damage portion to determine a size and a location of the damaged portion, determining a repair for the damaged portion based on the size and location of the damaged portion, and repairing the damaged portion with a composite repair kit. 
     Also disclosed is a composite repair kit for repairing the damaged portion of the composite fuselage or wing of the aircraft. The repair kit includes a plurality of nested sections and adhesive. Each nested section of the plurality of nested sections is a single-ply of composite material and the plurality of nested sections includes varying physical shapes to repair the damaged portion. Additionally, the nested sections, of the plurality of nested sections, are constructed to be stacked with each other to form different multi-ply composite structures. The adhesive is configured to attach a first nested section of the plurality of nested sections to a second nested section of the plurality of nested sections. 
     Other devices, apparatus, systems, methods, features and advantages of the invention will be or will become apparent to one with skill in the art upon examination of the following figures and detailed description. It is intended that all such additional systems, methods, features and advantages be included within this description, be within the scope of the invention, and be protected by the accompanying claims. 
    
    
     
       BRIEF DESCRIPTION OF THE FIGURES 
       The invention may be better understood by referring to the following figures. The components in the figures are not necessarily to scale, emphasis instead being placed upon illustrating the principles of the invention. In the figures, like reference numerals designate corresponding parts throughout the different views. 
         FIG. 1  is a flowchart of a known method for repairing large area damage on an aircraft utilizing composite materials. 
         FIG. 2  is a flowchart of an example of an implementation of a method for repairing a damaged portion of a composite fuselage or wing on an aircraft in accordance with the present disclosure. 
         FIG. 3A  is an exploded assembly view of an example of an implementation of repair of a damaged portion in accordance with the present disclosure. 
         FIG. 3B  is an assembly view of the example of the implementation of repair of the damaged portion, shown in  FIG. 3A , in accordance with the present disclosure. 
         FIG. 3C  is a zoomed-in exploded assembly view of the assembly view shown in  FIGS. 3A and 3B  in accordance with the present disclosure. 
         FIG. 4  is an exploded assembly view of the support stringer, first multi-ply composite hat-shaped section, and third multi-ply composite hat-shaped section, shown in  FIG. 4 , in accordance with the present disclosure. 
         FIG. 5  is a front assembly view of an example of an implementation of the nested combination, shown in  FIG. 4 , in accordance with the present disclosure. 
         FIG. 6A  is a prospective assembly view of an example of another implementation of a repair of a damaged portion of an aircraft in accordance with the present disclosure. 
         FIG. 6B  is an exploded prospective assembly view of the implementation of the repair of the damaged portion of an aircraft in accordance with the present disclosure. 
         FIG. 7A  is a prospective assembly view of an example of another implementation of the repair of a damaged portion of an aircraft in accordance with the present disclosure. 
         FIG. 7B  is an exploded assembly view of the implementation of the repair of the damaged portion, shown in  FIG. 7A , in accordance with the present disclosure. 
         FIG. 7C  is a cut-view of the exploded assembly of the implementation of the repair of the damaged portion, shown in  FIGS. 7A and 7B , in accordance with the present disclosure. 
         FIG. 8  is a perspective view of an example of an implementation of a single-ply hat-shaped section in accordance with the present disclosure. 
         FIG. 9  is a perspective view of an example of another implementation of a single-ply hat-shaped section in accordance with the present disclosure. 
         FIG. 10  is a perspective view of an example of yet another implementation of a single-ply hat-shaped section in accordance with the present disclosure. 
         FIG. 11A  is an assembly view of an example of an implementation of a multi-ply hat-shaped section in accordance with the present disclosure. 
         FIG. 11B  is an exploded assembly of the implementation of the multi-ply hat-shaped section, shown in  FIG. 11A , in accordance with the present disclosure. 
         FIG. 12A  is an assembly view of an example of another implementation of a multi-ply hat-shaped section in accordance with the present disclosure. 
         FIG. 12B  is an exploded assembly view of the implementation of a multi-ply hat-shaped section, shown in  FIG. 12A , in accordance with the present disclosure. 
         FIG. 13  is an exploded assembly view of an example of an implementation of a multi-ply structural element in accordance with the present disclosure. 
         FIG. 14  is a zoomed in assembly view of the implementation of the multi-ply structural element, shown in  FIG. 13 , in accordance with the present disclosure. 
         FIG. 15  is an exploded assembly view of an example of another implementation of a multi-ply structural element in accordance with the present disclosure. 
     
    
    
     DETAILED DESCRIPTION 
     A method for repairing a damaged portion of a composite fuselage or wing on an aircraft is disclosed. The method includes performing a non-destructive inspection (“NDI”) of the damage portion to determine a size and a location of the damaged portion, determining a repair for the damaged portion based on the size and location of the damaged portion, and repairing the damaged portion with a composite repair kit. 
     Also disclosed is a composite repair kit for repairing the damaged portion of the composite fuselage or wing of the aircraft. The repair kit includes a plurality of nested sections and adhesive. Each nested section of the plurality of nested sections is a single-ply of composite material and the plurality of nested sections includes varying physical shapes to repair the damaged portion. Additionally, the nested sections, of the plurality of nested sections, are constructed to be stacked with each other to form different multi-ply composite structures. The adhesive is configured to attach a first nested section of the plurality of nested sections to a second nested section of the plurality of nested sections. 
     In general, the composite repair kit allows for rapid, low cost fabrication of repair components for repairs on composite structures on aircraft and, in particular, for large area composite repairs. By utilizing the composite repair kit, stack-ups (i.e., stacking) of nested sections (that are pre-cured layers of composite material) may be utilized to bridge the boundary between existing structures in the aircraft and any new structures that are placed in the aircraft to repair the damaged portion of the fuselage or wing. Additionally, the stack-up of the nested sections may also be utilized to add thickness and stiffness to existing damaged structures within the aircraft rather than having to replace the damaged structure entirely. 
     In this example, the nested sections are pre-fabricated segmented composite laminate forms of pre-cured layers of material that may be rapidly and easily combined to produce splices between the repairs to the damaged portion and surrounding structure. Moreover, the nested sections may be utilized to produce skin doublers, stringers, and repair splice plates as will be described later. 
     For purposes of this disclosure, each of the nested sections may be a single-layered or a multi-layered composite laminate structure (also referred to, interchangeably, as a composite laminate form, a composite material, or a combined material) that was constructed from one or more layers of material. As such, in this disclosure, each nested section may be constructed as a consolidation of one or more (for example up to six) layers of fiber and matrix composite lamina that is bonded together with adhesive and cured to form a “single-ply” of material prior to being included in the composite repair kit. While it is generally understood that the term “ply” and layer may be interchangeably used in the art, for purpose of ease of description in this disclosure, the term “ply” will be limited to describing a resulting layer of combined material that has been cured from one or more component layers of material that were first bonded and then cured into the resulting combined material, which is herein referred to as simply a “ply of material” even though the ply of material may include the one or more component layers of material that were bonded and cured together. Since these nested sections are fabricated prior and then provided to the composite repair kit, the resulting nested sections in the composite repair kit will appear to be structures that have a single layer of composite material that are configured to be retrieved from the composite repair kit and then stacked together to form the multi-layer structures that may later be bonded together and cured by an end-user (of the composite repair kit) to form the multi-layered structures. As such, from the perspective of the end-user, each of the nested sections will appear to have a single layer of composite material even though that single layer may include multiple layers of component layers of material that form the single layer of composite material. Therefore, in this disclosure, a nested section of a single layer of composite material will herein be referred to as a “single-ply” section and nested sections that are combined into combined multi-layered nested sections will herein be referred to as a “multi-ply” section, since from the perspective of the end-user the plies will be based on the number of nested sections that were retrieved from the composite repair kit and combined not the actual number of layers of fiber and matrix composite lamina that were originally utilized to produce the nested sections prior to being supplied in the composite repair kit. 
     Specifically, in  FIG. 2 , a flowchart is shown of an example of an implementation of a method  200  for repairing a damaged portion of a composite fuselage or wing on an aircraft in accordance with the present disclosure. The method  200  starts  202  when a large area of damage occurs on the aircraft causing a damaged portion of the composite fuselage or wing on the aircraft. The method  200  then includes performing  204  a non-destructive inspection (“NDI”) of the damage portion to determine a size and location of the damaged portion. A repair procedure and design are then determined  206  for the damaged portion based on the size and location of the damaged portion, the nested sections are retrieved  208  from the composite repair kit, modified  210  if necessary, and the damaged portion is repaired  212  with the modified nested sections from the composite repair kit. The method then ends  214 . 
     In this example, determining a repair (i.e., a repair procedure and design) for repairing the damaged portion based on the size and location of the damaged portion may optionally include utilizing a commercial aviation services department (of the manufacturer of the aircraft) to determine what type of repair needs to be performed and then request that the support engineering department (of the manufacturer) analyzes the damaged portion and determine the repair procedure and design that is needed to properly repair the damaged portion. This repair procedure and design is determined, in part, based on the size and location of the damaged portion on the aircraft and on the types of nested sections in the composite repair kit. Moreover, determining the repair includes analyzing the damaged portion and designing a repair solution for the damaged portion that utilizes the composite repair kit, where the composite repair kit includes a plurality of nested sections and adhesive to attach different nested sections together to form varying multi-ply composite structures. 
     Once the repair procedure and design has been determined, local technicians near the aircraft are able to: retrieve the necessary nested sections from the composite repair kit; combine them into different multi-ply composite structures at the damaged portion of the aircraft utilizing the adhesive and other simple tools; cure the different multi-ply composite structures at the damaged portion (i.e., at the location of the damaged portion); and attach the different multi-ply composite structures at the damaged portion with either adhesive or bolts to complete the repair. The aircraft is then repaired and ready to enter service again. 
     As described earlier, the composite repair kit includes a plurality of nested sections and adhesive. Each of the nested sections are a single-ply of composite material and the composite repair kit includes different types of nested sections having varying physical shapes to repair different types of potential damaged portions of aircrafts. In general, the nested sections may include long narrow strips of single-ply material sheets of varying size and length, large area single-ply material sheets of varying size and length, curved flat single-ply material sheets of varying size and length, and single-ply hat-shaped sections of varying size and length. In this disclosure, each of the nested sections are autoclave cured consolidated elements (i.e., parts) that may be bonded together to create various structural elements. By producing the nested sections with an autoclave cure process prior to being supplied in the composite repair kit, the nested sections are higher quality composite elements that may be bonded together with adhesive at the location of the damaged portion of the aircraft with simple adhesive curing techniques. 
     In this example, the large area single-ply material sheets may be utilized to produce the repair skin or, when stacked and bonded together, the multi-ply composite repair skin. The narrow strips of single-ply material sheets may be utilized to produce a flat doubler, a multi-ply composite doubler (when stacked and bonded together), a single-ply splice plate, or a multi-ply composite splice plate when stacked and bonded together. The curve flat single-ply material sheets may also be utilized to produce the flat doubler, a multi-ply composite doubler (when stacked and bonded together), a single-ply splice plate, or a multi-ply composite splice plate when stacked and bonded together. Moreover, the narrow strips of single-ply material sheets may also be utilized to produce various structural elements such as stringers that may be, for example, I-shaped stringers, Z-shaped stringers, C-shaped stringers, or L-shaped stringers. Furthermore, with the aid of simple tools and relief cuts, the narrow strips of single-ply material sheets may be formed into complex contours since the narrow strips of single-ply material sheets may be designed to be strong but flexible. Moreover, the single-ply hat-shaped sections, or when stacked and bonded together, the multi-ply composite hat-shaped section, may be utilized to strengthen, repair stringers, replace stringers, splice stringers, bridge a stringer, create stringer doublers, or attached to other nested sections. 
     In aircraft, a stringer is generally a stiffening member that the skin of an aircraft is fastened to. In general, a stringer is: attached to a former (also known as frame) in a fuselage or to rib in a wing; a structural element that supports a section of the load carrying skin of the aircraft so as to prevent the skin from buckling under compression or shear loads; and primarily responsible for transferring the aerodynamic loads acting on the skin onto the frames or ribs of the aircraft. Based on the location and orientation of the stringer, the stringer may be referred to as a stringer or a longeron; however, for purposes of simplicity in this disclosure the term “stringer” will be utilized for both stringers and longerons. In general, stringers may be constructed of a strong and stiff material that is of acceptable weight and cost. Examples of the material utilized to construct stringers may include Aluminum 2024 T3, alloys of aluminum, steel, titanium, aluminum iron molybdenum zirconium, composite material such as carbon fiber and epoxy matrix resin, or other similar materials. 
     As an example, two or more nested sections may be trimmed with heavy duty scissors (e.g., compound scissors), bonded, and stacked to form multi-ply nested sections that may be utilized to repair stringers, replace stringers, splice stringers, and create stringer doublers. In this example, the nested sections are first trimmed and stacked (with adhesive), then cured in place (i.e., at the location of damaged portion) to ensure fit-up of the multi-ply nested section. Once the proper multi-ply nested section is created, the multi-ply nested section may be bolted or bonded in place. In general, by utilizing this approach, the resulting stepped sections in the damaged portion allow for good load paths. 
     In  FIG. 3A , an exploded assembly view of an example of an implementation of repair  300  of a damaged portion  302  is shown in accordance with the present disclosure. In this example, the damaged portion  302  includes a damaged section of a skin  304  of a composite fuselage or wing and a damaged section of a support stringer  306 . The repair  300  of the damaged portion  302  includes placing at least one repair skin  308  over the damaged portion  302 . The at least one repair skin  308  is a one or more nested section that is a one or more large area single-ply material sheet. In this example, the at least one repair skin  308  is shown placed internally over the damaged portion  302 . The at least one repair skin  308  may be a multi-ply composite repair skin, where the multi-ply composite repair skin includes a stacked up of a plurality of repair skins to form the multi-ply composite repair skin, where the plurality of repair skins are bonded together with adhesive. The repair  300  also includes one or more flat doublers  310  constructed of one or more single-ply material sheets placed around a periphery of the at least one repair skin  308  in a staggered and stacked manner to match a thickness of the at least one repair skin  308  to a thickness of the damaged section of the skin  304  of the composite fuselage or wing within the damaged portion  302 . In general, the flat doublers  310  are pre-cured laminates that include, for example, one to six plies thick of composite material. Usually, the flat doublers  310  will include two to three plies of material for ease of contouring and trimming. The one or more flat doublers  310  is at least one nested section that is a narrow strip of a single-ply material sheet. Moreover, the repair  300  also includes at least a single-ply splice plate  312  that is at least one nested section that is also a narrow strip of a single-ply material sheet  330  acting as a single-ply splice plate, where the at least a single-ply splice plate  312  is configured to splice (i.e., join) the at least one repair skin  308  to the skin  304  of the composite fuselage or wing. In this example, the one or more flat doublers  310  may be bonded to the skin  304  of the composite fuselage or wing and the at least a single splice plate  312  is attached to both the one or more flat doublers  310  and at least one repair skin  308  with either adhesive (i.e., bonded) or by mechanical means such as, for example, bolts. The repair  300  may also include a first multi-ply composite hat-shaped section  314  acting as a splice stringer, a second multi-ply composite hat-shaped section  316  acting a stringer doubler, and a third multi-ply composite hat-shaped section  318  acting as a restoration stringer. In this example, the first multi-ply composite hat-shaped section  314  may include two single-ply hat-shaped sections  320  and  322  bonded with adhesive, the second multi-ply composite hat-shaped section  316  may include a two single-ply hat-shaped sections  324  and  326  bonded with adhesive, and the third multi-ply composite hat-shaped section  318  may include a four single-ply hat-shaped sections  328 ,  330 ,  332 , and  334  bonded with adhesive. The single-ply hat-shaped sections  320 ,  322 ,  324 ,  326 ,  328 ,  330 ,  332 , and  334  are nested sections of the composite repair kit. 
     Turning to  FIG. 3B , an assembled assembly view of the example of the implementation of repair  300  of the damaged portion  302  is shown in accordance with the present disclosure. While not shown for purposes of ease of illustration, the single-ply hat-shaped sections  320 ,  322 ,  324 ,  326 ,  328 ,  330 ,  332 , and  334  may optionally be nested in a stepped fashion to allow for good load paths along the combined support stringer  306 , first multi-ply composite hat-shaped section  314 , and third multi-ply composite hat-shaped section  318 . Moreover, the single-ply hat-shaped sections  320 ,  322 ,  324 ,  326 ,  328 ,  330 ,  332 , and  334  may also be trimmed prior to being stacked up so as to produce the designed stack-up, such as, for example a stack-up that is nested in a stepped fashion. 
     In  FIG. 3C , a zoomed-in exploded assembly view is shown of the assembly view shown in  FIGS. 3A and 3B  in accordance with the present disclosure. In this example, each single-ply hat-shaped section  320 ,  322 ,  324 ,  326 ,  328 ,  330 ,  332 , or  334  is shown to be themselves optionally multi-layer structures as shown by a layer assembly  340 . As described earlier, for purposes of this disclosure the term “single-ply” means that the structure (i.e., a hat-shaped section or flat section) of the nested section is a single layer of composite material that was previously fabricated prior to being included in the composite repair kit and that may be stacked up and nested with other similar types of pre-cured structures to form a multi-layer structure that will be referred to a being “multi-ply.” As such, the term single-ply as used herein is not limited to the actual number of layers of material that were originally utilized (i.e., in the pre-cure process of fabrication) to produce the actual single-ply structure. As such, in this example, each single-ply hat-shaped section  320 ,  322 ,  324 ,  326 ,  328 ,  330 ,  332 , or  334  may be constructed of a single layer of material (if it has suitable material and mechanical properties) or it may include between two to six, or more, layers of material shown in the layer assembly  340  as layers P 1    342 , P 2    344 , through P 3    346 . Again, as described earlier, these layers are fabricated (including being cured in the fabrication process) prior to being provided to the composite repair kit to form the “single-ply” section that may then be combined with other “single-ply” sections to form “multi-ply” sections, where the term “ply.” in this disclosure, refers (from the perspective of the end-user) to the resulting composite layer of material that was previously fabricated and then provided to the composite repair kit. As such, the term “ply” refers to the layers utilized and combined by the end-user of the composite repair kit and is not limited to the actual number of layers of material utilized to produce a given nested section (in the original fabrication process prior to being provided to the composite repair kit) that is utilized by the end-user. In this example, each single-ply hat-shaped section  320 ,  322 ,  324 ,  326 ,  328 ,  330 ,  332 , or  334  may be fabricated (prior to being provided to the composite repair kit) as a consolidation of one to six layer (i.e., plies) of a fiber and matrix composite lamina such as, for example, a BMS 8-276 material (produced by The Boeing Company of Chicago, Ill.) and 6K-70-PW (a pre-impregnated “pre-preg” carbon fabric). 
     In  FIG. 4 , an exploded assembly view of the support stringer  306 , first multi-ply composite hat-shaped section  314 , and third multi-ply composite hat-shaped section  318  are shown in accordance with the present disclosure. In this example, all of the single-ply hat-shaped sections  320 ,  322 ,  324 ,  326 ,  328 ,  330 ,  332 , and  334  are shown as being bonded together with adhesive  400  placed between the single-ply hat-shaped sections  320 ,  322 ,  328 ,  330 ,  332 , and  334 . The combination (i.e., the nested combination  404 ) of all the single-ply hat-shaped sections  320 ,  322 ,  328 ,  330 ,  332 , and  334  are then cured together at the damaged portion  302  utilizing known portable heating and curing techniques. The nested combination  404  is cured in place to ensure proper fit-up on the support stringer  306  and second multi-ply composite hat-shaped section  316  and/or at least one repair skin  308 . In this example, the adhesive  400  may be a film adhesive. In this example, a release film may be placed between the bottom  406  of the nested combination  404  at the bottom single-ply hat-shaped section  334  and the top surface  408  of the support stringer  306 . The release film allows for removal and trimming after cure. Once removed and trimmed, the nested combination  404  may be stacked on top of the support stringer  306  and either bonded in place or bolted. It is appreciated by those of ordinary skill in the art that the single-ply hat-shaped sections  320 ,  322 ,  328 ,  330 ,  332 , and  334  are not drawn to scale in  FIG. 4  for the purpose of illustration and that the relative lengths of the individual single-ply hat-shaped sections  320 ,  322 ,  328 ,  330 ,  332 , and  334  may vary or may be as shown in  FIGS. 3A and 3B . 
     In  FIG. 5 , a front assembly view of an example of an implementation of the nested combination  404  is shown in accordance with the present disclosure. In this example in order to enable section use flexibility, some of the single-ply hat-shaped sections may be split at the top to allow for proper nesting. In general, the composite repair kit may include nested sections that are single-ply hat-shaped sections of varying sizes; however, in some situations it is necessary to attempt to nest two single-ply hat-shaped sections of the same size. As an example, a first split  500  is shown in the single-ply hat-shaped section  328  of the third multi-ply composite hat-shaped section  318  and a second split  502  is shown in the single-ply hat-shaped section  322  of the second multi-ply composite hat-shaped section  316 . In this example, the single-ply hat-shaped sections  322  and  328  may be the same shape and size of the single-ply hat-shaped sections  330  such that the first split  500  and the second split  502  allow both the single-ply hat-shaped sections  322  and  328  to fit and be nested between the single-ply hat-shaped sections  320  and  330 . In this example, to ensure acceptable bonding surfaces in the field, a peal-ply may be added to each surface of the single-ply hat-shaped sections  320 ,  322 ,  328 ,  330 ,  332 , and  334  and removed prior to bonding. 
       FIG. 6A  is a prospective assembly view of an example of another implementation of a repair  600  of a damaged portion  602  of an aircraft in accordance with the present disclosure. In this example, the repair  600  includes an at least one repair skin  604 , a first multi-ply composite hat-shaped section  606  acting as a splice stringer, a second multi-ply composite hat-shaped section  608  acting a stringer doubler, and a third multi-ply composite hat-shaped section  610  acting as a restoration stringer. In this example, the first multi-ply composite hat-shaped section  606  and the second multi-ply composite hat-shaped section  608  are located on a support stringer  612  and the first multi-ply composite hat-shaped section  606  and third multi-ply composite hat-shaped section  610  are located on the at least one repair skin  604 . Moreover, the repair  600  also includes a first multi-ply composite doubler  614 , a second multi-ply composite doubler  616 , a first multi-ply composite splice plate  618 , and a second multi-ply composite splice plate  620 . The first multi-ply composite doubler  614  and second multi-ply composite doubler  616 , as discussed earlier, are constructed from a plurality of flat doublers stacked up in a staggered fashion for load introduction, which are bonded together with adhesive. The bottoms of the first multi-ply composite doubler  614  and second multi-ply composite doubler  616  are then bonded to the surface of a skin  622  of a composite fuselage or wing. In this example, the first multi-ply composite doubler  614  is a curved structure as is the first multi-ply composite splice plate  618 . Both the first multi-ply composite doubler  614  and second multi-ply composite doubler  616  is placed around the periphery  624  of the at least one repair skin  604  in a staggered and stacked manner to match a thickness of the at least one repair skin  604  to a thickness of the damaged section of the skin  622  of the composite fuselage or wing within the damaged portion  602 . The first multi-ply composite doubler  614  and second multi-ply composite doubler  616  may be bonded together in a staggered fashion. 
     In this example, it is appreciated by those of ordinary skill in the art that the first multi-ply composite doubler  614  and second multi-ply composite doubler  616  and flat doublers in general may have direction specific properties (i.e., modulus) to allow for specific direction and orientation related uses. As seen in  FIG. 6A , the first multi-ply composite doubler  614  and second multi-ply composite doubler  616  are flat and curved to properly match the periphery  624  of the at least one repair skin  604 . As such, some of the multi-ply composite doublers and flat doublers are flat and round (i.e., curved) in different directions so as to be able to properly clock/orient the multi-ply composite doublers or flat doublers in the proper position to repair the damaged portion  602 . 
     The first multi-ply composite splice plate  618  and the second multi-ply composite splice plate  620  and placed on top of both the first multi-ply composite doubler  614  and second multi-ply composite doubler  616  and below the first multi-ply composite hat-shaped section  606 . In this example, the first multi-ply composite splice plate  618  is also a curved structure. The first multi-ply composite splice plate  618  and the second multi-ply composite splice plate  620  may be bonded together in a staggered fashion and either bonded or fastened by bolts to the first multi-ply composite doubler  614 , second multi-ply composite doubler  616 , and the at least one repair skin  604 . 
     Turning to  FIG. 6B , an exploded prospective assembly view is shown of the implementation of the repair  600  of the damaged portion  602  of an aircraft in accordance with the present disclosure. In this example, the first multi-ply composite hat-shaped section  606  includes two single-ply hat-shaped sections  606   a  and  606   b  bonded together with adhesive the second multi-ply composite hat-shaped section  608  includes four single-ply hat-shaped sections  608   a ,  608   b ,  608   c , and  608   d  bonded together with adhesive, and third multi-ply composite hat-shaped section  610  includes two single-ply hat-shaped sections  610   a  and  610   b  bonded together with adhesive. Similarly, the first multi-ply composite doubler  614  includes four flat doublers  614   a ,  614   b ,  614   c , and  614   d  bonded together with adhesive, second multi-ply composite doubler  616  includes four flat doublers  616   a ,  616   b ,  616   c , and  616   d  bonded together with adhesive, first multi-ply composite splice plate  618  includes four flat doublers  618   a ,  618   b ,  618   c , and  618   d  bonded together with adhesive, and second multi-ply composite splice plate  620  includes four flat doublers  620   a ,  620   b ,  620   c , and  620   d  bonded together with adhesive. 
     In  FIG. 7A , a prospective assembly view is shown of an example of another implementation of the repair  700  of a damaged portion  702  of an aircraft in accordance with the present disclosure. In this example, a multi-ply composite repair skin  704  is shown placed in the inside of the damaged portion  702 . The multi-ply composite repair skin  704  is held in place by a plurality of multi-ply composite splice plates  706  and  708  either via bonding adhesive or attachment bolts. The plurality of multi-ply composite splice plates  706  and  708  are stacked up on and attached to (via adhesive bonding) to a plurality of multi-ply composite doublers  710  and  712 . As described earlier, the plurality of multi-ply composite doublers  710  and  712  are located at the periphery of the multi-ply composite repair skin  704  and are attached to the skin  714  of a composite fuselage or wing having a plurality of support stringers  716 ,  718 ,  720 , and  722 . The repair  700  also include a plurality of multi-ply composite hat-shaped sections  724 ,  726 ,  728 , and  730  acting as splice stringers and a plurality of multi-ply composite hat-shaped sections  732  and  734  acting as restoration stringers. The repair may also include a plurality of multi-ply composite hat-shaped sections  736 ,  738 ,  740 , and  742  acting a stringer doublers and the plurality of multi-ply composite hat-shaped sections  724 ,  726 ,  728 , and  730  may have a staggered trim  744 ,  746 ,  748 , and  750  at the interface of the multi-ply composite hat-shaped sections  736 ,  738 ,  740 , and  742  and the plurality of multi-ply composite hat-shaped sections  732  and  734 . 
     In  FIG. 7B , an exploded assembly view is shown of the implementation of the repair  700  of the damaged portion  702  in accordance with the present disclosure. For the purpose of illustration simplicity the plurality of multi-ply composite hat-shaped sections  724 ,  726 ,  728 ,  730 ,  732  and  734  are not shown and the support stringers  716  and  718  are shown as damaged at the damaged portion  702 . In this example, the plurality of multi-ply composite doublers  710  and  712  are shown to include flat doublers  710   a ,  710   b ,  712   a , and  712   b , respectively. The multi-ply composite repair skin  704  include single-ply repair skins  704   a ,  704   b ,  704   c ,  704   d ,  704   e , and  704   f . The plurality of multi-ply composite splice plates  706  and  708  include single-ply splice plates  706   a ,  706   b ,  706   c ,  706   d ,  708   a ,  708   b ,  708   c , and  708   d , respectively, and the plurality of multi-ply composite doublers  710 ,  712 ,  750 , and  752  include flat doublers  710   a ,  710   b ,  712   a ,  712   b ,  752   a ,  752   b ,  754   a , and  754   b , respectively. Turning to  FIG. 7C , an exploded assembly cut-view is shown of the implementation of the repair  700  of the damaged portion  702  in accordance with the present disclosure. The cut-view is along plane AA′  756 . In this example, the plurality of multi-ply composite hat-shaped sections  724 ,  728 ,  732  and  734  are shown. 
     In  FIG. 8 , a perspective view is shown of an example of an implementation of a single-ply hat-shaped section  800  in accordance with the present disclosure. The single-ply hat-shaped section  800  includes a top surface  802 , a first side surface  804 , a second side surface  806 , a first bottom surface  808 , and a second bottom surface  810 . 
     In  FIG. 9 , a perspective view of an example of another implementation of a single-ply hat-shaped section  900  in accordance with the present disclosure. In this example, the single-ply hat-shaped section  900  also includes a top surface  902 , a first side surface  904 , a second side surface  906 , a first bottom surface  908 , and a second bottom surface  910 . The first bottom surface  908  includes a first portion  908   a  that has a first width  912  and a second portion  908   b  that has a second width  914  and the second bottom surface  910  includes a first portion  910   a  that also has the first width  912  and a second portion  910   b  that has a second width  914 . In this example, the second width  914  is greater than the first width  912 . 
     The single-ply hat-shaped section  900  may be a standard nested section from the composite repair kit or an end-user modified structure that has been trimmed to produce trimmed edges  916  and  918  of the second portions  908   b  and  910   b  between the first portions  908   a  and  910   a  and second portions  908   b  and  910   b , respectively. A plurality of single-ply hat-shaped sections, similar to the example single-ply hat-shaped section  900 , may be stacked up to produce a multi-ply hat-shaped section that has the shorter first length  912  of the combined first and second bottom surfaces. 
     Turning to  FIG. 10 , a perspective view of an example of yet another implementation of a single-ply hat-shaped section  1000  is shown in accordance with the present disclosure. In this example, the single-ply hat-shaped section  1000  also includes a top surface  1002 , a first side surface  1004 , a second side surface  1006 , a first bottom surface  1008 , and a second bottom surface  1010 . The first side surface  1004  has a first length  1012  (i.e., the length of the first side surface  1004 ) and includes a first portion  1004   a  that has a first width  1004   a   1  and a second portion  1004   b  that has a second width  1004   b   1 . In this example, the first width  1004   a   1  in the first portion  1004   a  is greater than the second width  1004   b   1  in the second portion  1004   b . Moreover, the first and second bottom surfaces  1008  and  1010  have a second length  1014  (i.e., the length of both the first and second bottom surfaces  1008 ) that is less than the first length  1012 . 
     The single-ply hat-shaped section  1000  may be a standard nested section from the composite repair kit or an end-user modified structure that has been trimmed to produce a trimmed edge  1016  of the second portion  1004   b  and the shorter second length  1014  of the first bottom surface  1008  and second bottom surface  1010 . As described earlier, a plurality of single-ply hat-shaped sections (similar to the example single-ply hat-shaped section  1000 ) may be stacked up to produce a multi-ply hat-shaped section that has the shorter second length  1014  of the combined first and second bottom surfaces. 
     In  FIG. 11A , an assembly view is shown of an example of an implementation of multi-ply hat-shaped section  1100  in accordance with the present disclosure. In this example, the multi-ply hat-shaped section  1100  includes a first portion  1102  and second portion  1104 . Both the first and second portions  1102  and  1104  include a plurality of single-ply hat-shaped sections that each include a top surface, a first side surface, a second side surface, a first bottom surface, and a second bottom surface; however, the bottom single-ply hat-shaped section  1106  of the second portion  1104  includes a first and second bottom surfaces  1108  and  1110  that are similar to the ones described in the example shown in  FIG. 9 . 
     In  FIG. 11B , an exploded assembly view is shown of the implementation of the multi-ply hat-shaped section  1100  in accordance with the present disclosure. The first portion  1102  of the multi-ply hat-shaped section  1100  includes a plurality of single-ply hat-shaped sections  1102   a ,  1102   b ,  1102   c ,  1102   d ,  1102   e , and  1102   f  and the second portion  1104  includes a plurality of single-ply hat-shaped sections that has the bottom single-ply hat-shaped section  1106  and single-ply hat-shaped sections  1104   a ,  1104   b ,  1104   c ,  1104   d , and  1104   e.    
     Turing to  FIG. 12A , an assembly view is shown of an example of another implementation of a multi-ply hat-shaped section  1200  in accordance with the present disclosure. In this example, the multi-ply hat-shaped section  1200  includes a plurality of single-ply hat-shaped sections  1200   a ,  1200   b ,  1200   c ,  1200   d ,  1200   e , and  1200   f  stacked up in a staggered manner and nested on top of each other. In  FIG. 12B , an exploded assembly view is shown of the implementation of a multi-ply hat-shaped section  1200  in accordance with the present disclosure. 
     As described earlier, the multi-ply hat-shaped sections  1100  and  1200  are formed by bonding the individual single-ply hat-shaped sections  1102   a ,  1102   b ,  1102   c ,  1102   d ,  1102   e ,  1102   f ,  1104   a ,  1104   b ,  1104   c ,  1104   d ,  1104   e ,  1106 ,  1200   a ,  1200   b ,  1200   c ,  1200   d ,  1200   e , and  1200   f , respectively. Similar to the example shown in  FIG. 5 , some of the individual single-ply hat-shaped sections  1102   a ,  1102   b ,  1102   c ,  1102   d ,  1102   e ,  1102   f ,  1104   a ,  1104   b ,  1104   c ,  1104   d ,  1104   e ,  1106 ,  1200   a ,  1200   b ,  1200   c ,  1200   d ,  1200   e , and  1200   f  may include splits (not shown) in the top surfaces to allow single-ply hat-shaped sections of the same shape and size to fit and be nested between varying layers of single-ply hat-shaped sections. As before, to ensure acceptable bonding surfaces in the field, a peal-ply may be added to each surface of the single-ply hat-shaped sections  1102   a ,  1102   b ,  1102   c ,  1102   d ,  1102   e ,  1102   f ,  1104   a ,  1104   b ,  1104   c ,  1104   d ,  1104   e ,  1106 ,  1200   a ,  1200   b ,  1200   c ,  1200   d ,  1200   e , and  1200   f  and removed prior to bonding. 
     In  FIG. 13 , an exploded assembly view is shown of an example of an implementation of a multi-ply structural element  1300  in accordance with the present disclosure. In this example, the multi-ply structural element  1300  may be a “Z-shaped” element because it will have a Z-shape when bonded and cured. In this example, the multi-ply structural element  1300  includes a first portion  1302  and second portion  1304  that include a plurality of single-ply nested sections  1302   a ,  1302   b ,  1302   c ,  1302   d ,  1304   a ,  1304   b ,  1304   c , and  1304   d , respectively. The single-ply nested sections  1302   a ,  1302   b ,  1302   c ,  1302   d ,  1304   a ,  1304   b ,  1304   c , and  1304   d  may be long and narrow strips that are bent into an “L-shape,” where the orientation of the L-shape is in a first direction for the single-ply nested sections  1302   a ,  1302   b ,  1302   c , and  1302   d  of the first portion  1302  and in an opposite direction for the single-ply nested sections  1304   a ,  1304   b ,  1304   c , and  1304   d  of the second portion  1304 . When bonded together with adhesive, the combined structure forms the multi-ply structural element  1300  having a Z-shape. In addition, an additional single-ply nested section  1306  may be bonded between the first portion  1302  and the second portion  1304  at the single-ply nested sections  1302   a  and  1304   a  to add thickness, strength, or both. Moreover, in order to form the multi-ply structural element  1300  that has complex contours, relief cuts  1308  may be cut into the plurality of single-ply nested sections  1302   a ,  1302   b ,  1302   c ,  1302   d ,  1304   a ,  1304   b ,  1304   c , and  1304   d , respectively. It is appreciated that for the purposes of ease of illustration, the relief cuts  1308  are shown only on the single-ply nested section  1302   d ; however, if present the relief cuts  1308  will also be within the other single-ply nested sections  1302   a ,  1302   b ,  1302   c ,  1304   a ,  1304   b ,  1304   c , and  1304   d . In  FIG. 14 , a zoomed in assembly view of the implementation of the multi-ply structural element  1300  is shown. In this example, a portion  1400  of the multi-ply structural element  1300  is shown with one relief cut  1402  of the plurality of relief cuts  1308  shown in  FIG. 13 . As before, the single-ply nested sections  1302   a ,  1302   b ,  1302   c ,  1302   d ,  1304   a ,  1304   b ,  1304   c ,  1304   d , and  1306  are bonded together with adhesive. 
     Turning to  FIG. 15 , an exploded assembly view of an example of another implementation of a multi-ply structural element  1500  is shown in accordance with the present disclosure. In this example, the multi-ply structural element  1500  forms an “L-shaped” element that may be similar to the first portion  1302  of the multi-ply structural element  1300  shown in  FIGS. 13 and 14 . The multi-ply structural element  1500  includes single-ply nested sections  1500   a ,  1500   b ,  1500   c , and  1500   d  that are long and narrow strips that are bent into an L-shape. Moreover, in this example, an additional single-ply nested section  1502  is bonded to the outer single-ply nested section  1500   d.    
     It is appreciated by those of ordinary skill in the art that while the composite repair kit is described as being utilized for repair and restoration of a damaged structure (i.e., the damaged portion), the composite repair kit may also be utilized create new original structures (e.g., curved stringers). The composite repair kit also enables simple tooling to create complex shaped parts. 
     It will be understood that various aspects or details of the invention may be changed without departing from the scope of the invention. It is not exhaustive and does not limit the claimed inventions to the precise form disclosed. Furthermore, the foregoing description is for the purpose of illustration only, and not for the purpose of limitation. Modifications and variations are possible in light of the above description or may be acquired from practicing the invention. The claims and their equivalents define the scope of the invention. 
     The flowchart and block diagrams in the different depicted example of implementations illustrate the architecture, functionality, and operation of some possible implementations of apparatuses and methods in an illustrative example. In this regard, each block in the flowchart or block diagrams may represent a module, a segment, a function, a portion of an operation or step, some combination thereof. 
     In some alternative examples of implementations, the function or functions noted in the blocks may occur out of the order noted in the figures. For example, in some cases, two blocks shown in succession may be executed substantially concurrently, or the blocks may sometimes be performed in the reverse order, depending upon the functionality involved. Also, other blocks may be added in addition to the illustrated blocks in a flowchart or block diagram. 
     The description of the different examples of implementations has been presented for purposes of illustration and description, and is not intended to be exhaustive or limited to the examples in the form disclosed. Many modifications and variations will be apparent to those of ordinary skill in the art. Further, different examples of implementations may provide different features as compared to other desirable examples. The example, or examples, selected are chosen and described in order to best explain the principles of the examples, the practical application, and to enable others of ordinary skill in the art to understand the disclosure for various examples with various modifications as are suited to the particular use contemplated.