초록 ▼

A randomly damaged area of a one-piece composite component of an aircraft is repaired according to a method herein. The damaged area covers skin and underlying stiffening substructure of the component. The method includes generating a design of a customized composite replacement panel for replacing

A randomly damaged area of a one-piece composite component of an aircraft is repaired according to a method herein. The damaged area covers skin and underlying stiffening substructure of the component. The method includes generating a design of a customized composite replacement panel for replacing the damaged area. The design includes replacement skin and underlying co-cured replacement stiffening substructure. The method further includes fabricating the composite replacement panel according to the design.

대표청구항 ▼

1. A method of repairing a randomly damaged area of a one-piece composite component of an aircraft, the damaged area covering skin and underlying stiffening substructure of the component, the method comprising: generating a design of a customized composite replacement panel for replacing the damaged

1. A method of repairing a randomly damaged area of a one-piece composite component of an aircraft, the damaged area covering skin and underlying stiffening substructure of the component, the method comprising: generating a design of a customized composite replacement panel for replacing the damaged area, the design including replacement skin and underlying integrated replacement stiffening substructure; andfabricating the composite replacement panel according to the design. 2. The method of claim 1, wherein the component is a one-piece composite barrel section of a fuselage. 3. The method of claim 1, wherein the replacement panel is designed to match contour of the damaged area. 4. The method of claim 1, wherein the damaged area is at least 3 feet×3 feet. 5. The method of claim 1, wherein generating the design includes generating a panel detail definition for the replacement panel and also a detail definition for elements for fastening the replacement panel to the component. 6. The method of claim 5, wherein the fastening element detail definition defines splice doublers for attaching the replacement panel to the component. 7. The method of claim 5, wherein the design specifies a one-piece replacement panel. 8. The method of claim 5, wherein the design specifies skin and integrated stiffeners, and also at least one excised element. 9. The method of claim 1, further comprising cutting damaged material from the damaged area, thereby leaving an opening in the component; and installing the fabricated replacement panel in the opening, including mechanically fastening the replacement panel to the component. 10. The method of claim 1, further comprising repairing small damaged areas via a bonded repair process. 11. The method of claim 1, wherein fabricating the replacement panel includes fabricating a custom layup mandrel tool and using the fabricated tool for layup and curing of the replacement panel. 12. The method of claim 11, wherein the mandrel tool is fabricated in a fabrication cell and used in the same fabrication cell to fabricate the replacement panel. 13. The method of claim 12, wherein the fabrication cell includes a dirty section for performing dirty operations on the mandrel tool during fabrication and on the panel after the panel has been cured; and a clean section for composite layup of the replacement panel on the mandrel tool. 14. The method of claim 13, further comprising using an end effector positioning system for performing the clean and dirty operations; and moving the end effector positioning system and the mandrel tool between the clean and dirty sections. 15. The method of claim 14, wherein the end effector positioning system uses a plurality of interchangeable end effectors to perform the clean and dirty operations. 16. The method of claim 13, wherein the dirty operations in the dirty section include trimming and machining of the mandrel tool and the cured panel; and wherein clean operations in the clean section include panel layup and material cutting. 17. The method of claim 13, further comprising performing nondestructive inspection of the replacement panel after curing. 18. The method of claim 13, further comprising a curing section, adjacent the clean section, for curing. 19. The method of claim 18, wherein: the mandrel tool is built and machined in the dirty section;composite material for the replacement panel is laid up on the mandrel tool in the clean section;the composite material for the panel is cured in the curing section; andthe cured panel is machined in the dirty section. 20. The method of claim 12, further comprising using the mandrel tool to fabricate composite splice doublers for fastening the replacement panel to the component. 21. The method of claim 11, wherein fabricating the replacement panel further includes mounting a face sheet of the mandrel tool to a rotary support and counterbalancing the face sheet; using the rotary support to rotate the mandrel tool; and depositing fabric on the face sheet as the mandrel tool is being rotated. 22. The method of claim 21, wherein the rotary support includes a spindle; and wherein the counterbalancing includes attaching weights with spacers to the spindle; and verifying balance conditions are satisfied to ensure rotational equilibrium within capabilities of a machine that is used to deposit the fabric on the face sheet. 23. The method of claim 22, wherein the spindle has an adjustable length; and wherein mounting the mandrel tool includes varying length of the spindle to accommodate length of the mandrel tool. 24. The method of claim 1, wherein generating the design of the replacement panel includes identifying width of tape originally used in the damaged area; and applying a set of rules governing material laydown. 25. The method of claim 24, wherein the rules relate to deviations and defects from laying down tape at a given width. 26. The method of claim 24, wherein the rules also identify a plurality of fabrication cells that have capability to fabricate the replacement panel at different tape widths. 27. The method of claim 26, wherein the rules identify those fabrication cells that achieve a best balance between (1) laydown machine configuration and tape width; (2) engineering requirements for composite laminate balance and symmetry, (3) structural perFormance, (4) weight of the replacement panel; and (5) speed of manufacturing the replacement panel. 28. A method of repairing a damaged one-piece composite fuselage barrel section of an aircraft fuselage, the method comprising: generating a design of a customized composite replacement panel for replacing a large damaged area of the barrel section, the designed panel including skin and integrated stiffening substructure;fabricating the replacement panel according to the design;fabricating fasteners for the replacement panel;cutting damaged material from the damaged area, thereby leaving an opening in the barrel section; andinstalling the fabricated panel in the opening, including using the fasteners to mechanically fasten the fabricated panel to the barrel section. 29. The method of claim 28, further comprising performing non-destructive inspection of the aircraft fuselage to determine location and extent of the damaged area. 30. The method of claim 29, wherein the non-destructive inspection is initiated upon receipt of a report indicating visible damage to the fuselage or a triggering action that might result in damage.