FIELD REPAIR KITS FOR STRUCTURAL PANEL

A kit for repairing an opening formed in a composite panel having a laminate skin arranged on opposing sides of a first core material is provided including a plug and a patch. The plug is formed from a second core material and has a cured filler material arranged within at least a portion of the second core material. The patch is formed from a plurality of cured first plies arranged in a stacked orientation. The plug is configured to be received within the opening of the composite panel and the patch is configured to cover the opening of the composite panel.

BACKGROUND OF THE INVENTION

Exemplary embodiments of the invention generally relate to a rotary wing aircraft, and more particularly, to a composite structural panel of a rotary wing aircraft.

Due to the favorable strength and weight characteristics of composite materials, the use of composites in various industries continues to expand. In aircraft manufacturing, the increasing use of composite material and composite structural assemblies is leading to significant improvements in fuel economy and substantial reduction of operating costs and atmospheric emissions. Currently, composite panels are used in the structural airframe of a rotary wing aircraft. These composite panels typically utilize a honeycomb core material having cured fiberglass or prepreg composite skins bonded thereto.

The composite panels in the airframe are commonly damaged during operation of the aircraft. Local damage may be repaired using a time consuming process that involves first curing a new piece of core material into the composite panel, and then secondly, curing a plurality of plies to patch the composite skin. Because the core material and composite skin are cured separately, the time required to repair the composite panel, and therefore the time that the rotary wing aircraft is non-operational, typically exceeds two days.

BRIEF DESCRIPTION OF THE INVENTION

According to one embodiment of the invention, a kit for repairing an opening formed in a composite panel having a laminate skin arranged on opposing sides of a first core material is provided including a plug and a patch. The plug is formed from a second core material and has a cured filler material arranged within at least a portion of the second core material. The patch is formed from a plurality of cured first plies arranged in a stacked orientation. The plug is configured to be received within the opening of the composite panel and the patch is configured to cover the opening of the composite panel.

According to another embodiment of the invention, a method of repairing a damaged composite panel of a rotary wing aircraft is provided including removing the damaged portion of laminate skin and an adjacent first core material of the composite panel to form an opening. A pre-formed plug having a bonding material applied to a portion thereof is inserted within the opening. The pre-formed plug includes a second core material having a cured filler material arranged within the second core material. A pre-formed patch having a bonding material applied to a portion thereof is installed over the opening so that a portion of the patch overlaps the adjacent laminate skin. The bonding material of the patch and the plug is cured to couple the patch and plug to the composite panel.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1Aschematically illustrates a rotary-wing aircraft10having a main rotor system12. The aircraft10includes an airframe14having an extending tail16which mounts a tail rotor system18, such as an anti-torque system for example. The main rotor assembly12is driven about an axis of rotation A through a main gearbox (illustrated schematically at T) by one or more engines E. The main rotor system12includes a plurality of rotor blades22mounted to a rotor hub20. Although a particular helicopter configuration is illustrated and described in the disclosed non-limiting embodiment, other configurations and/or machines, such as high speed compound rotary wing aircraft with supplemental translational thrust systems, dual contra-rotating coaxial rotor system aircraft, turboprops, tilt-rotors, and tilt-wing aircraft, will also benefit from the present invention.

Referring now toFIGS. 1B and 1C, the airframe14, may include a multitude of frame members24and a multitude of beam members26which support an aircraft outer skin28formed from a plurality of composite panels30(FIG. 2). The frame members24and beam members26may be arranged in a generally rectilinear pattern, however, any arrangement may be used as the composite panels30provide the rigidity necessary to reduce or eliminate the necessity of stringers. The composite panels30may be utilized in maintenance walkway areas, such as aircraft upper surfaces (FIG. 1B), high impact areas such as aircraft undersurfaces (FIG. 1C), as well as other areas such as wheel wells, floors, and steps of a rotary wing aircraft10.

Referring now toFIG. 2, a cross-section of a portion of a composite panel30is illustrated in more detail. The composite panel30may include a flange edge32configured to mechanically attach, such as with fasteners34for example, to the frame members24and/or beam members26of the airframe14. Each composite panel30includes multiple layers bonded together. The panel30includes a first laminate skin36aand a second, similar laminate skin36barranged adjacent opposing sides of a core material38. Each laminate skin36includes multiple plies, for example having graphite, fiberglass, or carbon fibers. The core material38may be a low density honeycomb core, such as the Kevlar® core material manufactured by DuPont™ Advanced Fibers Systems of Richmond, Va., USA, or may include an advanced core material such as K-COR™ or X-COR™ manufactured by Albany Engineered Composites of Mansfield, Mass., USA.

A damaged composite panel30can be repaired, rather than replaced, by removing the locally damaged portion of the laminate skin36and the adjacent portion of damaged core material38to form one or more openings40in the composite panel30(FIG. 3). Referring now toFIGS. 4 and 5, a pre-formed kit50including a plug52and patch54may be used to repair each of the openings40created in the composite panel30. In embodiments where the composite panel30has multiple openings40of different sizes, multiple repair kits50having plugs52and patches54of various sizes and shapes may be applied.

The pre-formed plug52includes a material, for example a low density honeycomb core material that can be substantially identical to the core material38of the composite panel30. In one embodiment, the material of the plug52has been pre-densified by curing the plug52after the openings (not shown) in the material of the plug52are filled with a filler material, such as EPOCAST® epoxy for example. The size and shape of the plug52is similar to at least one opening40formed in the core material38, and therefore to the portion of the core material38removed from the composite panel30. Although a plug52having a generally rectangular cross-section is illustrated, plugs52having any shape or cross-section are within the scope of the invention. When installed within the composite panel30, the plug52may be generally arranged within a plane (FIG. 5). However, in some embodiments, the plug52has a generally non-planar contour, such as a radius for example, that is generally complementary to the remainder of the composite panel30.

Similar to the laminate skin36, the patch54includes multiple stacked plies56cured to form a patch54having a thickness substantially equal to the thickness of the laminate skin36of the composite panel30. The material of the plies56used to form the patch54can be similar or identical to the plies used to form the laminate skin36. The size of the patch54can be substantially larger than the opening40formed in the skin over which the patch54is being applied. The patch54is configured to cover the opening40and overlap the adjacent laminate skin36about the entire periphery of the opening40. In one embodiment, the length and width of each ply56in the patch54varies such that the portion of each ply56that overlaps the adjacent laminate skin36differs dimensionally from that of another ply56in the patch54. As a result, the patch54can have a generally feathered or tapered edge58.

Referring now toFIG. 5andFIG. 6, a method100of repairing a damaged composite panel30using a pre-formed kit50is illustrated. In block102, a damaged portion of the laminate skin36and core material38are removed to form an opening40in the composite panel30. A plug52of the repair kit50, having a bonding material60applied about its outer periphery53, is inserted into an opening40in the composite panel30in block104. In block106, the patch54of the repair kit50, similarly having a bonding material60applied thereto, is arranged generally centrally over the plug52in an overlapping orientation with the adjacent laminate skin36. In one embodiment, the bonding material60is applied to the underside of the patch54, or the surface of the patch54configured to contact the plug52and the laminate skin36. The plug52and patch54are cured in block108to couple the plug52and patch54with the adjacent core material38and laminate skin36. To cure the patch54and plug52, a vacuum bagging film62is placed around the patch54and is secured to the surface of the laminate skin36. A port64extends through the bagging film62and is configured to connect to a vacuum pump, illustrated schematically at V, which applies a uniform pressure over the patch54.

The process of repairing a damaged portion of a composite panel30in an aircraft10is simplified by using the pre-fabricated plug52and patch54of the repair kit50. Because the plug52and patch54in the kit50are ready for assembly, only a structural epoxy adhesive and low pressure are required to cure the plug and patch to the composite panel. This can result in a significant time reduction for the repair of each composite panel, as compared with other repair processes.