Composite Repair Kits

Abstract

The present invention relates to a composite repair kit. A method for repairing a damaged portion of a composite fuselage or wing on an aircraft is disclosed. The method comprises: performing a non-destructive inspection ("NDI") on the damaged portion to determine the size and 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 using a composite repair kit.

Description

Technical Field

[0001] The present disclosure relates to methods and systems for repairing structures comprising composite materials, and in particular to methods and systems for performing such repairs on aircraft with limited resources and time. Background Art

[0002] The use of structures constructed of composite materials has increased significantly in many fields, including the aircraft industry. Reasons include the benefits of increased strength and rigidity, reduced weight, and reduced number of parts per structure. However, with the increase in the use of structures constructed of composite materials (e.g., composite structures), there is a need to properly repair any damage to these types of structures. In particular, for aircraft that employ composite structures, large-area composite repairs quickly become a significant support issue. As an example, although small-scale damage to an aircraft fuselage may require a miter repair, as the size of the damage increases, other methods are required so that: for large repairs, the integration of the repaired structure and the surrounding structure will require a significant amount of time to repair. In addition, composite structures often require extensive repair operations, which may ground the aircraft for a significant period of time (e.g., three weeks or more), thereby significantly increasing the support costs of the aircraft by taking the aircraft out of service.

[0003] In general, Figure 1 The current process for repairing large damage on aircraft using composite materials is described in . Figure 1 , a flow of a known method 100 is shown. When extensive damage occurs on an aircraft, the method 100 begins at 102. The known method 100 then includes a non-destructive inspection ("NDI") of the damage to determine the size and location of the damage 104. The information is then typically transmitted to the aircraft manufacturer's civil aviation service department, which determines what type of repair is required and then generates and sends a request (e.g., a repair definition) 106 to the manufacturer's support engineering department to analyze the damage and design a custom repair kit 108. The support engineering department then designs the custom repair kit 110 and sends a request to the manufacturer's manufacturing department to manufacture the custom repair kit. The manufacturing department then needs to commit tooling, materials, processing, labor, and manufacturing facility space 112 to manufacture 114 the custom repair kit, which may take, for example, three weeks or more to complete. Once the custom repair kit is completed, it is transported 116 to the site of the affected aircraft, where aircraft on ground ("AOG") personnel deliver and complete the repair to the aircraft 118. The method then ends 120.

[0004] This known method is a time consuming and expensive process. Therefore, there is a need for a system and method that is faster, more efficient and less expensive than current methods and allows for extensive damage repair of aircraft having composite structures. Summary of the invention

[0005] 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") on the damaged portion to determine the size and 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 using a composite repair kit.

[0006] A composite repair kit for repairing a damaged portion of a composite fuselage or wing of an aircraft is also disclosed. The composite repair kit includes a plurality of nests and an adhesive. Each of the plurality of nests is a single-ply composite material, and the plurality of nests include a physical shape that is altered to repair the damaged portion. In addition, some of the plurality of nests are configured to overlap each other to form different multi-ply composite structures. The adhesive is configured to attach a first nest of the plurality of nests to a second nest of the plurality of nests.

[0007] After reading the following drawings and detailed description, other devices, equipment, systems, methods, features and advantages of the present invention will be apparent to those skilled in the art. All these additional systems, methods, features and advantages are intended to be included in this specification, fall within the scope of the present invention and be protected by the appended claims. BRIEF DESCRIPTION OF THE DRAWINGS

[0008] The present invention will be better understood by referring to the accompanying drawings. The components in the accompanying drawings are not necessarily to scale, instead emphasis is placed on illustrating the principles of the present invention. In the accompanying drawings, the same reference numerals represent corresponding components throughout the different views.

[0009] Figure 1 is a flow chart of a known method for repairing extensive damage on an aircraft employing composite materials.

[0010] Figure 2 is a flow chart of an example implementation of a method for repairing a damaged portion of a composite fuselage or wing on an aircraft according to the present disclosure.

[0011] Figure 3A is an exploded assembly diagram of an implementation example of repairing a damaged part according to the present disclosure.

[0012] Figure 3B According to this disclosure Figure 3A An assembly diagram of an implementation example of repair of a damaged portion is shown in FIG.

[0013] Figure 3C According to this disclosure Figure 3A and Figure 3B An enlarged exploded assembly view of the assembly drawing shown in .

[0014] Figure 4 According to this disclosure Figure 4 An exploded assembly view of the support beam, the first multi-ply composite hat section, and the third multi-ply composite hat section shown in FIG.

[0015] Figure 5 According to this disclosure Figure 4 A front view assembly diagram of an implementation example of a nested combination shown in .

[0016] Fig. 6A It is a perspective assembly diagram of another implementation example of repairing a damaged part of an aircraft according to the present disclosure.

[0017] Figure 6B is an exploded perspective assembly diagram of an implementation of repairing a damaged portion of an aircraft according to the present disclosure.

[0018] Fig. 7A It is a perspective assembly diagram of another implementation example of repairing a damaged part of an aircraft according to the present disclosure.

[0019] Figure 7B According to this disclosure Fig. 7A Exploded assembly view of the implementation of repair of the damaged part shown in .

[0020] Figure 7C According to this disclosure Fig. 7A and Figure 7B Exploded cutaway assembly view of the implementation of repair of the damaged portion shown in FIG.

[0021] Figure 8 4 is a perspective view of an implementation example of a single-layer cap-shaped portion according to the present disclosure.

[0022] Fig. 9 4 is a perspective view of another implementation example of a single-layer cap-shaped portion according to the present disclosure.

[0023] Fig.10 4 is a perspective view of another implementation example of a single-layer cap-shaped portion according to the present disclosure.

[0024] Fig.11A is an assembly diagram of an implementation example of a multi-layer sheet hat-shaped portion according to the present disclosure.

[0025] Fig. 11B According to this disclosure Fig.11A Exploded assembly of the implementation of the multi-layer sheet hat-shaped portion shown in FIG.

[0026] Fig. 12A 1 is an assembly diagram of another implementation example of the multi-layered cap-shaped portion according to the present disclosure.

[0027] Fig. 12B According to this disclosure Fig. 12AAn exploded assembly view of an implementation of a multi-layer sheet hat-shaped portion is shown in FIG.

[0028] Fig.13 is an exploded assembly diagram of an implementation example of a multilayer sheet structure element according to the present disclosure.

[0029] Fig.14 According to this disclosure Fig.13 An enlarged assembly view of an implementation of the multilayer sheet structural element shown in FIG.

[0030] Fig.15 FIG. 4 is an exploded assembly diagram of another embodiment of a multilayer structural element according to the present disclosure. DETAILED DESCRIPTION

[0031] 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") on the damaged portion to determine the size and 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 using a composite repair kit.

[0032] A composite repair kit for repairing a damaged portion of a composite fuselage or wing of an aircraft is also disclosed. The composite repair kit includes a plurality of nests and an adhesive. Each of the plurality of nests is a single-ply composite material, and the plurality of nests include a physical shape that is altered to repair the damaged portion. In addition, some of the plurality of nests are configured to overlap each other to form different multi-ply composite structures. The adhesive is configured to attach a first nest of the plurality of nests to a second nest of the plurality of nests.

[0033] In general, the composite repair kit allows for rapid, low-cost manufacturing of repair parts for repairing composite structures on aircraft and particularly for large-area composite repairs. By using the composite repair kit, stacked (i.e., superimposed) nests (which are pre-cured layers of composite materials) are used to bridge the boundary between existing structures in the aircraft and any new structures placed in the aircraft to repair damaged portions of the fuselage or wings. In addition, the stacked nests are also used to increase the thickness and stiffness of existing damaged structures within the aircraft rather than having to replace the damaged structure in its entirety.

[0034] In this embodiment, the nested portion is in the form of a prefabricated segmented composite laminate of pre-cured layers of material that can be quickly and easily assembled to create a splice between the damaged portion and the repair of the surrounding structure. In addition, the nested portion can be used to create skin reinforcement panels, stringers and repair splices, which will be described later.

[0035] For purposes of this disclosure, each nest may be a single or multi-layer composite laminate structure (also interchangeably referred to as a composite laminate form, composite material, or composite material) comprised of one or more layers of material. To this end, in this disclosure, each nest may be constructed as a consolidation of one or more layers (e.g., up to six layers) of fiber and matrix composite layups that are bonded together with an adhesive and cured to form a "single ply" of material prior to inclusion in a composite repair kit. Although it is generally understood that the terms "ply" and "layer" may be used interchangeably in the art, in this disclosure, for ease of description purposes, the term "ply" will be limited to describing a final layer of a composite material that has been cured from one or more constituent material layers that are first bonded and then cured into a final composite material, which will be referred to herein simply as a "ply of material," even though the ply of material may include one or more constituent material layers bonded and cured together. Since these nested portions are first manufactured and then provided to the composite repair kit, the final nested portions in the composite repair kit will appear to have a structure having a single layer of composite material, which is configured to be retrieved from the composite repair kit and then stacked together to form a multi-layer structure, which can then be combined and cured by the end user (of the composite repair kit) to form a multi-layer structure. In this way, from the perspective of the end user, each nested portion will appear to have a single layer of composite material, even though the single layer may include multiple layers of constituent material layers that form the single layer of composite material. Therefore, in the present disclosure, nested portions of single layers of composite materials are referred to herein as "single-layer sheets" portions, and nested portions combined into combined multi-layered nested portions are referred to herein as "multi-layer sheets" portions, because from the perspective of the end user, the plies will be based on the number of nested portions retrieved and combined from the composite repair kit, rather than based on the actual number of layers of fiber and matrix composite laminates initially used to generate the nested portions before being supplied to the composite repair kit.

[0036] Specifically, in Figure 2 , a flow chart of an example implementation of a method 200 for repairing a damaged portion of a composite fuselage or wing on an aircraft according to the present disclosure is shown. The method 200 begins at 202 when extensive damage occurs on an aircraft resulting in a damaged portion of a composite fuselage or wing on the aircraft. The method 200 then includes performing 204 a non-destructive inspection ("NDI") of the damaged portion to determine the size and location of the damaged portion. A repair step and design is then determined 206 for the damaged portion based on the size and location of the damaged portion, a nest is retrieved 208 from a composite repair kit, modified 210 if necessary, and the damaged portion is repaired 212 with the modified nest from the composite repair kit. The method then ends 214.

[0037] In this embodiment, determining a repair (i.e., repair steps and design) for repairing the damaged portion based on the size and location of the damaged portion may optionally include utilizing a civil aviation service department (of the aircraft manufacturer) to determine what type of repair is needed and then requesting a support engineering department (of the manufacturer) to analyze the damaged portion and determine the repair steps and design required to appropriately repair the damaged portion. The repair steps and design are determined based in part on the size and location of the damaged portion on the aircraft and the type of nesting portion in the composite repair kit. In addition, determining the repair includes analyzing the damaged portion and designing a repair solution for the damaged portion using a composite repair kit, wherein the composite repair kit includes a plurality of nesting portions and an adhesive to attach different nesting portions together to form a modified multi-layer composite structure.

[0038] Once the repair steps and design have been determined, a local technician near the aircraft can: retrieve the required nesting parts from the composite repair kit; combine the nesting parts into different multi-layer composite structures at the damaged part of the aircraft using adhesives and other simple tools; cure the different multi-layer composite structures at the damaged part (i.e., at the location of the damaged part); and attach the different multi-layer composite structures at the damaged part with adhesives or bolts to complete the repair. The aircraft is then repaired and ready to be put into service again.

[0039] As previously described, the composite repair kit includes a plurality of nested sections and an adhesive. Each nested section is a single-ply composite material, and the composite repair kit includes different types of nested sections with varying physical shapes to repair different types of potentially damaged portions of an aircraft. In general, the nested sections may include: long narrow strips of single-ply material sheets with varying sizes and lengths; large-area single-ply material sheets with varying sizes and lengths; curved flat single-ply material sheets with varying sizes and lengths; and single-ply hat-shaped sections with varying sizes and lengths. In the present disclosure, each nested section is an autoclave-cured consolidated element (i.e., component) that can be combined together to generate various structural elements. By generating the nested sections through an autoclave curing process before being supplied to the composite repair kit, the nested sections are higher quality composite elements that are combined together with an adhesive at the location of the damaged portion of the aircraft through a simple adhesive curing technique.

[0040] In this embodiment, large area single-ply sheet material sheets can be used to generate repair skins, or when stacked and combined together, can be used to generate multi-ply composite repair skins. Long narrow strips of single-ply sheet material sheets can be used to generate flat reinforcement panels, multi-ply composite reinforcement panels (when stacked and combined together), single-ply spliced ​​panels, or when stacked and combined together to generate multi-ply composite spliced ​​panels. Curved flat single-ply sheet material sheets can also be used to generate flat reinforcement panels, multi-ply composite reinforcement panels (when stacked and combined together), single-ply spliced ​​panels, or when stacked and combined together to generate multi-ply composite spliced ​​panels. In addition, long narrow strips of single-ply sheet material sheets can also be used to generate various structural elements, such as, for example, stringers that can be I-shaped stringers, Z-shaped stringers, C-shaped stringers, or L-shaped stringers. In addition, with the help of simple tools and process cuts, the long narrow strips of single-ply sheet material sheets can be formed into complex contours because they can be designed to be strong but flexible. Furthermore, single ply hat sections or multi-ply composite hat sections when stacked and bonded together may be used to reinforce, repair stringers, replace stringers, splice stringers, bridge stringers, create stringer stiffeners, or attach to other nested sections.

[0041] In an aircraft, a stringer is typically a reinforcing member used for fastening on the skin of the aircraft. Typically, a stringer: is attached to a former (also called a frame) in the fuselage or to a rib in a wing; is a structural element that supports a portion of the load-bearing skin of the aircraft to prevent the skin from deforming under compression or shear loads; and is primarily responsible for transferring aerodynamic loads acting on the skin to the frame or rib of the aircraft. Based on the location and orientation of the stringer, the stringer may be referred to as a stringer or a longitudinal beam; however, for the purpose of simplicity in this disclosure, the term "stringer" will be used for both stringers and longitudinal beams. Typically, stringers may be constructed of a strong material that is acceptable in weight and cost. Embodiments of materials used to construct stringers may include aluminum 2024T3, aluminum alloys, steel, titanium, aluminum iron molybdenum zirconium, composite materials such as carbon fiber and epoxy matrix resin, or other similar materials.

[0042] As an example, two or more nests may be trimmed, joined and stacked by heavy-duty scissors (e.g., composite scissors) to form a multi-layer nest that can be used to repair stringers, replace stringers, splice stringers and generate stringer reinforcements. In this embodiment, the nest is first trimmed and stacked (with adhesive) and then cured in the appropriate position (i.e., at the location of the damaged portion) to ensure that the multi-layer nest is placed. Once a suitable multi-layer nest is generated, the multi-layer nest can be bolted or joined in the appropriate position. Overall, by adopting this method, the final step in the damaged portion allows a good load path.

[0043] exist Figure 3A, an exploded assembly view of an implementation example of a repair 300 of a damaged portion 302 according to the present disclosure is shown. In this embodiment, the damaged portion 302 includes a damaged portion of a skin 304 of a composite fuselage or wing and a damaged portion 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 one or more nested portions of one or more large-area single-ply sheet material sheets. In this embodiment, the at least one repair skin 308 is shown to be placed internally over the damaged portion 302. The at least one repair skin 308 can be a multi-ply composite repair skin, wherein the multi-ply composite repair skin includes a stack of multiple repair skins to form a multi-ply composite repair skin, wherein the multiple repair skins are bonded together with an adhesive. The repair 300 also includes one or more flat stiffening panels 310, which are composed of one or more single-ply sheet material sheets placed around the perimeter of the at least one repair skin 308 in a staggered and overlapping manner so that the thickness of the at least one repair skin 308 matches the thickness of the damaged portion of the composite fuselage or wing skin 304 within the damaged portion 302. Generally, the flat stiffening panels 310 are, for example, pre-cured laminates of composite material that include one to six plies thick. Generally, for ease of configuration and trimming, the flat stiffening panels 310 include two to three plies of sheet material. The one or more flat stiffening panels 310 are at least one nested portion of narrow strips of single-ply sheet material sheets. In addition, the repair 300 also includes at least one single-ply splice plate 312 that is at least one nested portion, which is also a narrow strip of single-ply material sheet 330 used as a single-ply splice plate, wherein the at least one single-ply splice plate 312 is configured to splice (i.e., join) the at least one repair skin 308 to the composite fuselage or wing skin 304. In this embodiment, the one or more flat stiffener panels 310 can be bonded to the composite fuselage or wing skin 304, and the at least one single-ply splice plate 312 is attached to both the one or more flat stiffener panels 310 and the at least one repair skin 308 with an adhesive (i.e., bonding) or by means of a mechanical mechanism (such as, for example, bolts). The repair 300 can also include a first multi-ply composite hat section 314 used as a splice stringer, a second multi-ply composite hat section 316 used as a stringer stiffener, and a third multi-ply composite hat section 318 used as a restoration stringer. In this embodiment, the first multi-layer sheet composite cap-shaped portion 314 may include two single-layer sheet cap-shaped portions 320 and 322 bonded together with an adhesive, the second multi-layer sheet composite cap-shaped portion 316 may include two single-layer sheet cap-shaped portions 324 and 326 bonded together with an adhesive, and the third multi-layer sheet composite cap-shaped portion 318 may include four single-layer sheet cap-shaped portions 328, 330, 332 and 334 bonded together with an adhesive.The single layer sheet hats 320, 322, 324, 326, 328, 330, 332 and 334 are nested portions of the composite repair kit.

[0044] Steering Figure 3B , shows an assembled assembly diagram of an example embodiment of a repair 300 of a damaged portion 302 according to the present disclosure. Although not shown for 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 manner to allow for a good load path along the combined support beam 306, the first multi-ply composite hat-shaped section 314, and the third multi-ply composite hat-shaped section 318. In addition, the single-ply hat-shaped sections 320, 322, 324, 326, 328, 330, 332, and 334 may also be trimmed prior to stacking to create a designed stack, such as, for example, a stack nested in a stepped manner.

[0045] exist Figure 3C In the figure, it is shown that according to the present disclosure Figure 3A and Figure 3B340 . In this embodiment, each single-ply cap 320, 322, 324, 326, 328, 330, 332, or 334 is shown as itself optionally a multi-layer structure as shown as layer assembly 340. As previously stated, for the purposes of this disclosure, the term "single-ply" means that the structure of the nested portion (i.e., the cap or flat portion) is a single layer of composite material that has been previously manufactured prior to inclusion in a composite repair kit and can be stacked and nested with other similar types of pre-cured structures to form a multi-layer structure (which will be referred to as a "multi-ply"). As such, the term single-ply, as used herein, is not limited to the actual number of layers of material that are initially used (i.e., during the pre-curing manufacturing process) to generate the actual single-ply structure. Thus, in this embodiment, each single-ply cap-shaped portion 320, 322, 324, 326, 328, 330, 332 or 334 may be composed of a single layer of material (if it has suitable material and mechanical properties), or may include between two to six or more layers of material (shown as layers P1 342, P2 344 to P3 346 in the layer assembly 340). Again as previously described, these layers are manufactured (including being cured during the manufacturing process) before being provided to the composite repair kit to form a "single-ply" portion, which can then be combined with other "single-ply" portions to form a "multi-ply" portion, where the term "ply" in this disclosure refers to the final composite material layer previously manufactured (from the perspective of the end user) and then provided to the composite repair kit. Thus, the term "ply" refers to the layers that are adopted and combined by the end user of the composite repair kit, and is not limited to the actual number of layers of material used to generate a given nested portion adopted by the end user (in the initial manufacturing process before being provided to the composite repair kit). In this embodiment, each single-ply cap portion 320, 322, 324, 326, 328, 330, 332 or 334 can be manufactured (before being provided to the composite repair kit) as a consolidation of one to six layers (i.e., plies) of fiber and matrix composite layups such as, for example, BMS 8-276 material (manufactured by The Boeing Company of Chicago, Illinois) and 6K-70-PW (pre-impregnated "prepreg" carbon fiber fabric).

[0046] exist Figure 4, an exploded assembly view of a support stringer 306, a first multi-ply composite hat 314, and a third multi-ply composite hat 318 according to the present disclosure is shown. In this embodiment, all of the single-ply hats 320, 322, 324, 326, 328, 330, 332, and 334 are shown to be bonded together by an adhesive 400 placed between the single-ply hats 320, 322, 328, 330, 332, and 334. The combination of all of the single-ply hats 320, 322, 328, 330, 332, and 334 (i.e., the nested combination 404) is then cured together at the damaged portion 302 using known portable heating and curing techniques. The nested combination 404 is cured in place to ensure proper placement on the support stringer 306 and the second multi-ply composite hat 316 and / or at least one repair skin 308. In this embodiment, the adhesive 400 may be an adhesive film. In this embodiment, a release film may be placed between the bottom 406 of the nested assembly 404 at the bottom single-layer cap 334 and the top surface 408 of the support beam 306. The release film allows for removal and trimming after curing. Once removed and trimmed, the nested assembly 404 can be stacked on top of the support beam 306 and bonded in place or screwed. It will be appreciated by those skilled in the art that in Figure 4 For the purpose of illustration, the single-layer sheet cap-shaped portions 320, 322, 328, 330, 332 and 334 are not drawn to scale, and the relative lengths of the individual single-layer sheet cap-shaped portions 320, 322, 328, 330, 332 and 334 may vary or may be as shown. Figure 3A and Figure 3B shown.

[0047] exist Figure 5, a front view assembly diagram of an implementation example of a nesting combination 404 according to the present disclosure is shown. In this embodiment, in order to allow flexibility in the use of the part, some single-layer sheet hat-shaped parts can be split at the top to allow proper nesting. In general, the composite repair kit can include nesting parts that are single-layer sheet hat-shaped parts with varying sizes; however, in some cases it is necessary to try to nest two single-layer sheet hat-shaped parts with the same size. As an example, a first crack 500 in the single-layer sheet hat-shaped part 328 of the third multi-layer sheet composite hat-shaped part 318 is shown, and a second crack 502 in the single-layer sheet hat-shaped part 322 of the second multi-layer sheet composite hat-shaped part 316 is shown. In this embodiment, the single-layer sheet hat-shaped parts 322 and 328 can have the same shape and size as the single-layer sheet hat-shaped part 330, so that the first crack 500 and the second crack 502 allow both the single-layer sheet hat-shaped parts 322 and 328 to be assembled and nested between the single-layer sheet hat-shaped parts 320 and 330. In this embodiment, to ensure an acceptable bonding surface in the field, peel plies may be added to each surface of the single ply hats 320, 322, 328, 330, 332, and 334 and removed prior to bonding.

[0048] Fig. 6A6 is a perspective assembly diagram of another embodiment of a repair 600 of a damaged portion 602 of an aircraft according to the present disclosure. In this embodiment, the repair 600 includes at least one repair skin 604, a first multi-layer composite hat-shaped portion 606 used as a splicing stringer, a second multi-layer composite hat-shaped portion 608 used as a stringer reinforcement plate, and a third multi-layer composite hat-shaped portion 610 used as a restoration stringer. In this embodiment, the first multi-layer composite hat-shaped portion 606 and the second multi-layer composite hat-shaped portion 608 are located on a supporting stringer 612, and the first multi-layer composite hat-shaped portion 606 and the third multi-layer composite hat-shaped portion 610 are located on the at least one repair skin 604. In addition, the repair 600 also includes a first multi-layer composite reinforcement plate 614, a second multi-layer composite reinforcement plate 616, a first multi-layer composite splicing plate 618, and a second multi-layer composite splicing plate 620. As previously described, the first multi-layer composite reinforcement panel 614 and the second multi-layer composite reinforcement panel 616 are composed of a plurality of flat reinforcement panels stacked in a staggered manner for load introduction and bonded together with an adhesive. The bottom of the first multi-layer composite reinforcement panel 614 and the second multi-layer composite reinforcement panel 616 are then bonded to the surface of the composite fuselage or wing skin 622. In this embodiment, the first multi-layer composite reinforcement panel 614 is a curved structure, as is the first multi-layer composite splice panel 618. The first multi-layer composite reinforcement panel 614 and the second multi-layer composite reinforcement panel 616 are both placed around the periphery 624 of the at least one repair skin 604 in a staggered and stacked manner so that the thickness of the at least one repair skin 604 matches the thickness of the damaged portion of the composite fuselage or wing skin 622 within the damaged portion 602. The first multi-layer composite reinforcement panel 614 and the second multi-layer composite reinforcement panel 616 can be bonded together in a staggered manner.

[0049] In this embodiment, those skilled in the art will appreciate that the first multi-layer composite reinforcement plate 614 and the second multi-layer composite reinforcement plate 616 and the flat reinforcement plate as a whole may have direction specificity (i.e., modulus) to allow for specific directional and orientation-related uses. Fig. 6A As seen in FIG. 6 , the first multi-ply composite stiffener 614 and the second multi-ply composite stiffener 616 are flat and curved to appropriately match the perimeter 624 of the at least one repair skin 604. In this way, some multi-ply composite stiffeners and flat stiffeners are flat and meander (i.e., curved) in different directions, so that the multi-ply composite stiffeners or flat stiffeners can be appropriately turned / oriented in the appropriate position to repair the damaged portion 602.

[0050] The first multi-ply composite splice plate 618 and the second multi-ply composite splice plate 620 are placed on top of the first multi-ply composite reinforcement plate 614 and the second multi-ply composite reinforcement plate 616 and below the first multi-ply composite hat-shaped portion 606. In this embodiment, 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 can be combined together in a staggered manner and combined or fastened to the first multi-ply composite reinforcement plate 614, the second multi-ply composite reinforcement plate 616 and at least one repair skin 604 by means of bolts.

[0051] Steering Figure 6B , shows an exploded perspective assembly view of an implementation of a repair 600 of a damaged portion 602 of an aircraft according to the present disclosure. In this embodiment, a first multi-ply composite hat-shaped portion 606 includes two single-ply hat-shaped portions 606a and 606b bonded together with an adhesive, a second multi-ply composite hat-shaped portion 608 includes four single-ply hat-shaped portions 608a, 608b, 608c, and 608d bonded together with an adhesive, and a third multi-ply composite hat-shaped portion 610 includes two single-ply hat-shaped portions 610a and 610b bonded together with an adhesive. Similarly, the first multi-layer composite reinforcement panel 614 includes four flat reinforcement panels 614a, 614b, 614c and 614d bonded together by adhesive, the second multi-layer composite reinforcement panel 616 includes four flat reinforcement panels 616a, 616b, 616c and 616d bonded together by adhesive, the first multi-layer composite splicing panel 618 includes four flat reinforcement panels 618a, 618b, 618c and 618d bonded together by adhesive, and the second multi-layer composite splicing panel 620 includes four flat reinforcement panels 620a, 620b, 620c and 620d bonded together by adhesive.

[0052] exist Fig. 7A, a perspective assembly view of another embodiment of a repair 700 of a damaged portion 702 of an aircraft according to the present disclosure is shown. In this embodiment, a multi-ply composite repair skin 704 is shown placed within the damaged portion 702. The multi-ply composite repair skin 704 is held in place by a plurality of multi-ply composite splice panels 706 and 708 with the aid of bonding adhesives or attachment bolts. The plurality of multi-ply composite splice panels 706 and 708 are stacked on and attached (with the aid of adhesive bonding) to a plurality of multi-ply composite reinforcement panels 710 and 712. As previously described, the plurality of multi-ply composite reinforcement panels 710 and 712 are located at the periphery of the multi-ply composite repair skin 704 and are attached to the skin 714 of the composite fuselage or wing having a plurality of support stringers 716, 718, 720, and 722. The repair 700 also includes a plurality of multi-ply composite hat-shaped sections 724, 726, 728, and 730 used as spliced ​​stringers and a plurality of multi-ply composite hat-shaped sections 732 and 734 used as restored stringers. The repair may also include a plurality of multi-ply composite hat-shaped sections 736, 738, 740, and 742 used as stringer reinforcement plates, and the plurality of multi-ply composite hat-shaped sections 724, 726, 728, and 730 may have staggered trimming sections 744, 746, 748, and 750 at the interfaces 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.

[0053] exist Figure 7B , an exploded assembly view of an implementation of a repair 700 of a damaged portion 702 according to the present disclosure is shown. For the purpose of simplicity of illustration, the plurality of multi-ply composite hat-shaped portions 724, 726, 728, 730, 732, and 734 are not shown, and the support beams 716, 718 are shown as being damaged at the damaged portion 702. In this embodiment, the plurality of multi-ply composite stiffeners 710 and 712 are shown to include flat stiffeners 710a, 710b, 712a, and 712b, respectively. The multi-ply composite repair skin 704 includes single-ply repair skins 704a, 704b, 704c, 704d, 704e, and 704f. The plurality of multi-ply composite splice plates 706 and 708 include single-ply splice plates 706a, 706b, 706c, 706d, 708a, 708b, 708c and 708d, respectively, and the plurality of multi-ply composite reinforcement plates 710, 712, 750 and 752 include flat reinforcement plates 710a, 710b, 712a, 712b, 752a, 752b, 754a and 754b, respectively. Figure 7C , shows an exploded assembly cutaway view of an implementation of a repair 700 of a damaged portion 702 according to the present disclosure. The cutaway view is along plane AA' 756. In this embodiment, a plurality of multi-ply composite hat portions 724, 728, 732, and 734 are shown.

[0054] exist Figure 8 , a perspective view of an implementation example of a single-layer cap-shaped portion 800 according to the present disclosure is shown. The single-layer cap-shaped portion 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.

[0055] exist Fig. 9 , a perspective view of another embodiment of a single-layer cap-shaped portion 900 according to the present disclosure is shown. In this embodiment, the single-layer cap-shaped portion 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 908a having a first width 912 and a second portion 908b having a second width 914, and the second bottom surface 910 includes a first portion 910a also having a first width 912 and a second portion 910b having a second width 914. In this embodiment, the second width 914 is greater than the first width 912.

[0056] The single-ply sheet-hat section 900 may be a standard nested section from a composite repair kit, or an end-user modified structure that has been trimmed to create trimmed edges 916 and 918 between first sections 908a and 910a and second sections 908b and 910b, respectively, of second sections 908b and 910b. Multiple single-ply sheet-hat sections similar to the embodiment single-ply sheet-hat section 900 may be stacked to create a multi-ply sheet-hat section having a shorter first length 912 combining a first bottom surface and a second bottom surface.

[0057] Steering Fig.10 , a perspective view of another embodiment of a single-layer cap-shaped portion 1000 according to the present disclosure is shown. In this embodiment, the single-layer cap-shaped portion 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 1004a having a first width 1004a1 and a second portion 1004b having a second width 1004b1. In this embodiment, the first width 1004a1 in the first portion 1004a is greater than the second width 1004b1 in the second portion 1004b. In addition, the first bottom surface 1008 and the second bottom surface 1010 have a second length 1014 (i.e., the length of both the first bottom surface 1008 and the second bottom surface) that is less than the first length 1012.

[0058] The single-layer sheet-hat section 1000 may be a standard nested section from a composite repair kit, or an end-user modified structure that has been trimmed to generate a first bottom surface 1008 and a shorter second length 1014 of the second bottom surface 1010 and a trimmed edge 1016 of the second portion 1004b. As previously described, multiple single-layer sheet-hat sections (similar to the embodiment single-layer sheet-hat section 1000) may be stacked to generate a multi-layer sheet-hat section having a shorter second length 1014 of a combined first bottom surface and a second bottom surface.

[0059] exist Fig.11A , an assembly diagram of an implementation example of a multilayer sheet cap-shaped portion 1100 according to the present disclosure is shown. In this embodiment, the multilayer sheet cap-shaped portion 1100 includes a first portion 1102 and a second portion 1104. The first portion 1102 and the second portion 1104 both include a plurality of single-layer sheet cap-shaped portions, each of which includes a top surface, a first side surface, a second side surface, a first bottom surface, and a second bottom surface; however, the bottom single-layer sheet cap-shaped portion 1106 of the second portion 1104 includes a similar Fig. 9 The first bottom surface 1108 and the second bottom surface 1110 are described in the embodiment shown in FIG.

[0060] exist Fig. 11B , an exploded assembly view of an implementation of a multilayer sheet hat 1100 according to the present disclosure is shown. The first portion 1102 of the multilayer sheet hat 1100 includes a plurality of single-layer sheet hats 1102a, 1102b, 1102c, 1102d, 1102e, and 1102f, and the second portion 1104 includes a plurality of single-layer sheet hats having the bottom single-layer sheet hat 1106 and single-layer sheet hats 1104a, 1104b, 1104c, 1104d, and 1104e.

[0061] Steering Fig. 12A , shows an assembly diagram of another embodiment of a multilayer sheet hat-shaped portion 1200 according to the present disclosure. In this embodiment, the multilayer sheet hat-shaped portion 1200 includes a plurality of single-layer sheet hat-shaped portions 1200a, 1200b, 1200c, 1200d, 1200e, and 1200f stacked in a staggered manner and nested on top of each other. Fig. 12B , an exploded assembly view of an implementation of a multi-layer sheet hat 1200 according to the present disclosure is shown.

[0062] As previously described, multi-layered sheet hats 1100 and 1200 are formed by combining the individual single-layered sheet hats 1102a, 1102b, 1102c, 1102d, 1102e, 1102f, 1104a, 1104b, 1104c, 1104d, 1104e, 1106, 1200a, 1200b, 1200c, 1200d, 1200e, and 1200f, respectively. Figure 5 In the embodiments shown in , some of the individual single-layer sheet hat portions 1102a, 1102b, 1102c, 1102d, 1102e, 1102f, 1104a, 1104b, 1104c, 1104d, 1104e, 1106, 1200a, 1200b, 1200c, 1200d, 1200e and 1200f may include cracks (not shown) in the top surface to allow single-layer sheet hat portions of the same shape and size to be assembled and nested between varying layers of the single-layer sheet hat portions. As before, to ensure an acceptable bonding surface in the field, peel plies may be added to the various surfaces of the single-ply cap portions 1102a, 1102b, 1102c, 1102d, 1102e, 1102f, 1104a, 1104b, 1104c, 1104d, 1104e, 1106, 1200a, 1200b, 1200c, 1200d, 1200e and 1200f and removed prior to bonding.

[0063] exist Fig.13, an exploded assembly view of an example implementation of a multi-layer sheet structure element 1300 according to the present disclosure is shown. In this embodiment, the multi-layer sheet structure element 1300 can be a "Z-shaped" element because it will have a Z-shape when combined and cured. In this embodiment, the multi-layer sheet structure element 1300 includes a first portion 1302 and a second portion 1304, which respectively include a plurality of single-layer sheet nests 1302a, 1302b, 1302c, 1302d, 1304a, 1304b, 1304c and 1304d. The single-ply nests 1302a, 1302b, 1302c, 1302d, 1304a, 1304b, 1304c, and 1304d may be long narrow strips bent into an "L-shape," wherein the L-shape is oriented in a first direction for the single-ply nests 1302a, 1302b, 1302c, and 1302d of the first portion 1302, and in an opposite direction for the single-ply nests 1304a, 1304b, 1304c, and 1304d of the second portion 1304. When bonded together with an adhesive, the combined structure forms a multi-ply structural element 1300 having a Z-shape. Additionally, additional single-ply nests 1306 may be bonded between the first portion 1302 and the second portion 1304 at the single-ply nests 1302a and 1304a to increase thickness, strength, or both. In addition, in order to form a multi-layer sheet structural element 1300 with a complex contour, process cuts 1308 can be cut in the plurality of single-layer sheet nests 1302a, 1302b, 1302c, 1302d, 1304a, 1304b, 1304c, and 1304d, respectively. It should be understood that for ease of illustration, the process cuts 1308 are only shown on the single-layer sheet nest 1302d; however, if present, the process cuts 1308 are also located in the other single-layer sheet nests 1302a, 1302b, 1302c, 1304a, 1304b, 1304c, and 1304d. Fig.14 , an enlarged assembly view of an implementation of a multilayer sheet structure element 1300 is shown. In this embodiment, a portion 1400 of the multilayer sheet structure element 1300 is shown having Fig.13 One process cut 1402 of the plurality of process cuts 1308 is shown in FIG. As before, the single-ply nests 1302a, 1302b, 1302c, 1302d, 1304a, 1304b, 1304c, 1304d, and 1306 are bonded together with an adhesive.

[0064] Steering Fig.15 , shows an exploded assembly diagram of another embodiment of a multilayer sheet structural element 1500 according to the present disclosure. In one embodiment, the multilayer sheet structural element 1500 may be formed similar to Fig.13 and Fig.14The "L-shaped" element of the first portion 1302 of the multi-layer sheet structural element 1300 shown in FIG. The multi-layer sheet structural element 1500 includes single-layer sheet nesting parts 1500a, 1500b, 1500c and 1500d which are long narrow strips bent into an L shape. In addition, in this embodiment, an additional single-layer sheet nesting part 1502 is bonded to the outer single-layer sheet nesting part 1500d.

[0065] It will be appreciated by those skilled in the art that although the composite repair kit is described as being used to repair and restore a damaged structure (i.e., a damaged portion), the composite repair kit can also be used to create a new original structure (e.g., a bent stringer). The composite repair kit also enables the creation of complex formed parts using simple tools.

[0066] It should be understood that various aspects or details of the present invention may be changed without departing from the scope of the present invention. It is not exclusive and does not limit the claimed invention to the exact form disclosed. In addition, the foregoing description is only for illustrative purposes, not for limiting purposes. Modifications and changes are feasible in view of the above description, or can be obtained from practicing the present invention. The claims and their equivalents define the scope of the present invention.

[0067] The flow charts and block diagrams in the different implementation examples described show the architecture, functionality and operation of some possible implementations of the device and method with illustrative embodiments. For this reason, each block in the flow chart and block diagram can represent a part of a module, segment, function, operation or step, or some combination thereof.

[0068] In some alternative implementation examples, one or more functions marked in the blocks may not appear in the order marked in the figure. For example, in some cases, two blocks shown in succession may be executed substantially simultaneously, or the blocks may sometimes be executed in reverse order, depending on the functionality involved. Moreover, in addition to the blocks shown in the flow chart or block diagram, other blocks may be added.

[0069] Furthermore, the present disclosure includes embodiments according to the following clauses:

[0070] Clause 1. A method for repairing a damaged portion of a composite fuselage or wing on an aircraft, the method comprising:

[0071] performing a non-destructive inspection ("NDI") on the damaged portion to determine the size and location of the damaged portion;

[0072] Determining repair of the damaged portion based on the size and location of the damaged portion; and

[0073] The damaged portion is repaired using a composite repair kit.

[0074] Clause 2. The method of clause 1, wherein determining the repair comprises:

[0075] analyzing the damaged portion; and

[0076] A repair scheme using the composite repair kit is designed for the damaged portion.

[0077] Clause 3. The method according to clause 2, wherein repairing the damaged portion with a composite repair kit comprises:

[0078] The damaged portion is repaired using a plurality of nested portions of the composite repair kit, wherein

[0079] Each nest of the plurality of nests is a single layer of composite material, and

[0080] The plurality of nested portions include a nested portion having a physical shape that changes to repair the damaged portion,

[0081] wherein the nested portions are configured to be stacked upon one another to form a modified multi-ply composite structure; and

[0082] An adhesive is applied at the location of the damaged portion on the aircraft to attach the first nest portion to the second nest portion.

[0083] Clause 4. The method of clause 3, wherein the plurality of nests are cured at the location of the damaged portion on the aircraft.

[0084] Clause 5. A method according to clause 3,

[0085] wherein the plurality of nested parts include at least one repair skin, a plurality of flat reinforcement panels and a plurality of single-layer splicing panels,

[0086] Wherein repairing the damaged part comprises:

[0087] placing the at least one repair skin internally over the damaged portion, wherein the damaged portion comprises a damaged portion of a skin of a composite fuselage or wing,

[0088] placing the plurality of flat reinforcement panels in a staggered and overlapping manner around the periphery of the at least one repair skin so that the thickness of the at least one repair skin matches the thickness of the skin of the composite fuselage or wing at the damaged portion; and

[0089] The plurality of single-ply splice plates are placed on the plurality of flat reinforcement plates and at least one repair skin in a staggered manner, wherein the splice plates join the plurality of flat reinforcement plates and the repair skin together.

[0090] Clause 6. A method according to clause 5,

[0091] wherein the at least one repair skin is a plurality of repair skins, and

[0092] Wherein placing the at least one repair skin over the damaged portion comprises:

[0093] applying an adhesive between the plurality of repair skins; and

[0094] The plurality of repair skins are stacked to form a multi-ply composite repair skin that is internally positioned over the damaged portion.

[0095] Clause 7. The method of Clause 6, further comprising: allowing the multi-ply composite repair skin cured at the location of the damaged portion to cure on the aircraft.

[0096] Clause 8. The method of clause 5, wherein placing the plurality of flat reinforcement panels around a perimeter of the at least one repair skin comprises:

[0097] applying an adhesive between the plurality of flat reinforcement sheets;

[0098] stacking the plurality of flat stiffening panels in a staggered manner to form a multi-ply composite stiffening panel, wherein stacking the plurality of flat stiffening panels further comprises matching the thickness of the at least one repair skin with the thickness of the composite fuselage or wing skin at the damaged portion; and

[0099] The multi-ply composite reinforced panel is cured on the aircraft at the location of the damaged portion.

[0100] Clause 9. The method of clause 5, wherein placing the plurality of single-ply splice panels over the plurality of flat stiffener panels and at least one repair skin comprises:

[0101] applying adhesive between the plurality of single-ply splice panels;

[0102] stacking the plurality of single-layer spliced ​​panels to form a multi-layer composite spliced ​​panel; and

[0103] The multi-ply composite splice panel is cured on the aircraft at the location of the damaged portion.

[0104] Clause 10. A method according to clause 3,

[0105] wherein the plurality of nested portions include a plurality of single-layer cap-shaped portions,

[0106] Wherein repairing the damaged part comprises:

[0107] applying adhesive between the plurality of single-layer sheet hat-shaped portions;

[0108] stacking the plurality of single-layered sheet hat-shaped portions to form a multi-layered sheet composite hat-shaped portion; and

[0109] The multi-ply composite hat section is cured on the aircraft at the location of the damaged portion.

[0110] Clause 11. The method of Clause 10, further comprising: applying a release film at the bottom single-ply sheet hat portion before curing the multi-ply sheet composite hat portion.

[0111] Clause 12. A method according to clause 10,

[0112] Each of the single-layer cap-shaped portions comprises a top surface, a first side surface, a second side surface, a first bottom surface, and a second bottom surface.

[0113] wherein the first bottom surface includes a first portion having a first width and a second portion having a second width, and

[0114] wherein one of the single-layer sheet hat-shaped portions is a wide-bottomed single-layer sheet hat-shaped portion having the second width greater than the first width,

[0115] Wherein stacking the plurality of single-layer sheet hat-shaped portions to form a multi-layer sheet composite hat-shaped portion comprises: stacking the plurality of single-layer sheet hat-shaped portions so that the wide-bottom single-layer sheet hat-shaped portion is the bottom of the multi-layer sheet composite hat-shaped portion.

[0116] Clause 13. The method according to clause 3, wherein repairing the damaged portion with a composite repair kit comprises:

[0117] applying adhesive between at least three of the nested portions; and

[0118] The at least three nested portions are combined to form a structural element for use in the aircraft.

[0119] Clause 14. A composite repair kit for repairing a damaged portion of a composite fuselage or wing, the composite repair kit comprising:

[0120] Multiple nested sections, where

[0121] Each nest of the plurality of nests is a single layer of composite material, and

[0122] The plurality of nested portions include a physical shape that changes to repair the damaged portion,

[0123] wherein certain nested portions of the plurality of nested portions are configured to overlap one another to form a modified multi-ply composite structure;

[0124] An adhesive is provided, wherein the adhesive attaches a first nest of the plurality of nests to a second nest of the plurality of nests.

[0125] Clause 15. The composite repair kit of Clause 14, wherein the plurality of nests includes a nest that is a repair skin configured to be placed internally over the damaged portion, wherein the damaged portion comprises a damaged portion of skin of the composite fuselage or wing.

[0126] Clause 16. A composite repair kit according to Clause 15, wherein the plurality of nested portions further include a plurality of flat reinforcement panels, the plurality of flat reinforcement panels being configured to be placed in a staggered manner around the periphery of the repair skin and stacked so that the thickness of the repair skin matches the thickness of the skin of the composite fuselage or wing at the damaged portion.

[0127] Clause 17. A composite repair kit according to Clause 16, wherein the plurality of nested portions further include a plurality of single-ply splicing plates, the plurality of single-ply splicing plates being configured to be positioned in a staggered manner above the plurality of flat reinforcement plates and the repair skin, wherein the splicing plates join the plurality of flat reinforcement plates and the repair skin together.

[0128] Clause 18. The composite repair kit of Clause 17, wherein the plurality of nested sections further comprises a plurality of single-ply sheet hat sections configured to bridge edges of stringers positioned above the repair skin.

[0129] Clause 19. The composite repair kit of Clause 14, wherein the plurality of nested sections comprises a plurality of single-ply sheet hat sections configured to bridge edges of stringers located proximate to a damaged portion of the composite fuselage or wing.

[0130] Clause 20. A composite repair kit according to clause 19,

[0131] Each of the single-layer cap-shaped portions comprises a top surface, a first side surface, a second side surface, a first bottom surface, and a second bottom surface.

[0132] wherein the single-layer cap-shaped portion has a first length,

[0133] wherein the top surface has the first length, and

[0134] wherein the first bottom surface and the second bottom surface have a second length, and

[0135] The first length is greater than the second length.

[0136] Clause 21. A composite repair kit according to clause 19,

[0137] Each of the single-layer cap-shaped portions comprises a top surface, a first side surface, a second side surface, a first bottom surface, and a second bottom surface.

[0138] wherein the first bottom surface includes a first portion having a first width and a second portion having a second width, and

[0139] The second width is greater than the first width.

[0140] The description of different implementation examples has been provided for the purpose of illustration and description, and is not intended to be exclusive or limited to the embodiments of the disclosed forms. Many modifications and variations are apparent to those skilled in the art. In addition, different implementation examples may provide different features compared to other desired embodiments. The selected one or more embodiments are selected and described to best explain the principles of the embodiments, practical applications, and to enable those skilled in the art to understand the present disclosure for various embodiments with various modifications suitable for the specific purposes under consideration.

Claims

1. A method (200) for repairing a damaged portion (302, 602, 702) of a composite fuselage or wing on an aircraft, the method (200) comprising: performing (204) a non-destructive inspection ("NDI") on the damaged portion (302, 602, 702) to determine the size and location of the damaged portion (302, 602, 702); Determining (206) a repair (600) for the damaged portion (302, 602, 702) based on the size and location of the damaged portion (302, 602, 702); and Repairing (212) the damaged portion (302, 602, 702) using a composite repair kit; Wherein, repairing (212) the damaged portion (302, 602, 702) using the composite repair kit comprises: The damaged portion (302, 602, 702) is repaired (600) using a plurality of nested portions (320, 322, 324, 326, 328, 330, 332, 334) of the composite repair kit, wherein: Each nested portion (320, 322, 324, 326, 328, 330, 332, 334) of the plurality of nested portions is a single-layer composite material, The plurality of nests include a nest having a changed physical shape to repair (600) the damaged portion (302, 602, 702), and The plurality of nested portions (320, 322, 324, 326, 328, 330, 332, 334) are configured to be stacked upon one another to form a nested stack of altered multi-ply composite structures (606, 608, 610, 614, 616, 618, 620); and applying an adhesive (400) at the location of the damaged portion (302, 602, 702) on the aircraft to attach the first nest to the second nest, wherein the plurality of nested portions include at least one repair skin, a plurality of flat reinforcement panels (310, 620a, 620b, 620c, 620d, 710a, 710b, 712a, 712b, 752a, 752b, 754a, 754b) and a plurality of single-layer splice panels (706a, 706b, 706c, 706d, 708a, 708b, 708c, 708d), Wherein repairing (212) the damaged portion (302, 602, 702) comprises: placing the at least one repair skin internally over the damaged portion (302, 602, 702), wherein the damaged portion (302, 602, 702) comprises a damaged portion (302, 602, 702) of the skin of the composite fuselage or wing, placing the plurality of flat reinforcement panels (310, 620a, 620b, 620c, 620d, 710a, 710b, 712a, 712b, 752a, 752b, 754a, 754b) around a perimeter (624) of the at least one repair skin in a staggered and overlapping manner so that the thickness of the at least one repair skin matches the thickness of the skin of the composite fuselage or wing at the damaged portion (302, 602, 702); and The plurality of single-layer splicing plates are placed in a staggered manner on the plurality of flat reinforcing plates (310, 620a, 620b, 620c, 620d, 710a, 710b, 712a, 712b, 752a, 752b, 754a, 754b) and at least one repair skin, wherein the splicing plates (706a, 706b, 706c, 706d, 708a, 708b, 708c, 708d) join the plurality of flat reinforcing plates (310, 620a, 620b, 620c, 620d, 710a, 710b, 712a, 712b, 752a, 752b, 754a, 754b) and the repair skin together.

2. The method (200) of claim 1, wherein determining (206) a repair (600) comprises: analyzing the damaged portion (302, 602, 702); as well as A repair scheme using the composite repair kit is designed for the damaged portion (302, 602, 702).

3. The method (200) according to claim 1, wherein the at least one repair skin (604) is a plurality (624) of repair skins, and Wherein placing the at least one repair skin (604) over the damaged portion (302, 602, 702) comprises: applying adhesive between the plurality of repair skins (400); as well as stacking the plurality of repair skins to form a multi-ply composite repair skin (704) placed internally on the damaged portion (302, 602, 702), and The method further includes curing the multi-ply composite repair skin (704) cured at the location of the damaged portion (302, 602, 702) on the aircraft.

4. The method (200) of claim 3, wherein placing the plurality of flat reinforcement panels (310, 620a, 620b, 620c, 620d, 710a, 710b, 712a, 712b, 752a, 752b, 754a, 754b) around a perimeter (624) of the at least one repair skin (604) comprises: applying an adhesive (400) between the plurality of flat reinforcement plates (310, 620a, 620b, 620c, 620d, 710a, 710b, 712a, 712b, 752a, 752b, 754a, 754b); stacking the plurality of flat stiffening panels (310, 620a, 620b, 620c, 620d, 710a, 710b, 712a, 712b, 752a, 752b, 754a, 754b) in a staggered manner to form a multi-ply composite stiffening panel (710, 712, 750, 752), wherein stacking the plurality of flat stiffening panels (310, 620a, 620b, 620c, 620d, 710a, 710b, 712a, 712b, 752a, 752b, 754a, 754b) further comprises matching the thickness of the at least one repair skin (604) with the thickness of the skin of the composite fuselage or wing at the damaged portion (302, 602, 702); as well as The multi-ply composite reinforced panel (710, 712, 750, 752) is cured on the aircraft at the location of the damaged portion (302, 602, 702).

5. The method (200) of claim 1, wherein placing the plurality of single-ply splice panels (706a, 706b, 706c, 706d, 708a, 708b, 708c, 708d) over the plurality of flat stiffener panels (310, 620a, 620b, 620c, 620d, 710a, 710b, 712a, 712b, 752a, 752b, 754a, 754b) and at least one repair skin (604) comprises: Applying adhesive (400) between the plurality of single-ply splice panels (706a, 706b, 706c, 706d, 708a, 708b, 708c, 708d); stacking the plurality of single-ply splice panels (706a, 706b, 706c, 706d, 708a, 708b, 708c, 708d) to form a multi-ply composite splice panel (618, 620); and The multi-ply composite splice panel (618, 620) is cured on the aircraft at the location of the damaged portion (302, 602, 702).

6. The method (200) according to claim 1, wherein the plurality of nested portions further comprises a plurality of single-layer cap-shaped portions (608a, 608b, 608c, 608d), Wherein repairing the damaged portion (302, 602, 702) includes: applying adhesive (400) between the plurality of single-layer sheet hat-shaped portions (608a, 608b, 608c, 608d); stacking the plurality of single-layered hat-shaped portions (608a, 608b, 608c, 608d) to form a multi-layered composite hat-shaped portion (320, 322, 324, 326, 328, 330, 332, 334, 610a, 610b); and The multi-ply composite hat portion (320, 322, 324, 326, 328, 330, 332, 334, 610a, 610b) is cured on the aircraft at the location of the damaged portion (302, 602, 702).

7. The method (200) according to claim 6, Each of the single-layer cap-shaped portions (608a, 608b, 608c, 608d) comprises a top surface, a first side surface, a second side surface, a first bottom surface, and a second bottom surface. wherein the first bottom surface includes a first portion having a first width and a second portion having a second width, and wherein one of the single-layer sheet hat-shaped portions (608a, 608b, 608c, 608d) is a wide-bottomed single-layer sheet hat-shaped portion having the second width greater than the first width, The method of stacking the plurality of single-layer cap-shaped portions (608a, 608b, 608c, 608d) to form a multi-layer composite cap-shaped portion (320, 322, 324, 326, 328, 330, 332, 334, 610a, 610b) comprises: The plurality of single-layer sheet hat-shaped portions (608a, 608b, 608c, 608d) are stacked so that the wide-bottom single-layer sheet hat-shaped portion is the bottom of the multi-layer sheet composite hat-shaped portion.

8. The method according to claim 1, further comprising: attaching each nest in the nested stack to another nest in the nested stack by an adhesive; placing the nested stack over the damaged portion; curing the nested stack while the nested stack is on the damaged portion; as well as The cured nested stack is attached to the damaged portion.

9. The method according to claim 8, further comprising: applying a release film between a bottom surface of the nested stack and a top surface of the damaged portion so as to be separated from the damaged portion after the nested stack is cured; as well as After separating the nested stack, the nested stack is trimmed based on the size and shape of the damaged portion.

Images