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Rapid Manufacturing of Tooling for On-Aircraft Composite Scarf Repairs

Description:

OBJECTIVE: Adhesively bonded on-aircraft composite scarf repairs require a lengthy, labor intensive, temporary tooling fabrication process. The goal is to develop and demonstrate rapid manufacturing of splash tooling that can withstand cures up to 350-450°F. Elimination of unnecessary mold forming steps and associated labor/time is desired such that mold fabrication time is reduced by 75%. 

DESCRIPTION: Current procedures for structural repair of polymer matrix composite (PMC) moderately contoured outer mold line (OML) surfaces involve a rather small but time-consuming master mold, or splash tooling, fabrication process. In a general repair process, the damaged solid epoxy or bismaleimide PMC material is removed at the required scarf ratio. The OML is then used for fabrication of a male mold using a plaster-like material. After approximately eight hours, the male form fabrication process is complete. Then, the male form is used to generate a female form using the same plaster material. After another eight hours of fabrication and cure, the female temporary form, which now mimics the OML, is now ready for single-use. The epoxy or bismaleimide carbon structural repair prepreg plies are laid-up, debulked, and cured in an autoclave per the prescribed cure cycle. Depending on the material to be replaced and repaired (i.e. epoxy or bismaleimide), the final cure temperatures will be in excess of 350°F to 450°F, and the final patch will experience consolidation pressure up to 100 psi to suppress interlaminar void formation. The pre-cured, or “hard,” scarf patch is now ready for structural adhesive bonding on aircraft. What is desired is a rapid manufacturing methodology that can eliminate the multi-step temporary tool (both male and female) fabrication process and withstand the harsh autoclave environments used specifically for epoxy and bismaleimide cures. Current additive manufacturing and 3D printing technologies appear promising for temporary tooling fabrication. However, even the highest performance unfilled polymer powders and filaments, such as polyetherimide (PEI), polyphenylene sulfide (PPS), and polyetheretherketone (PEEK), cannot withstand the final cure/postcure temperatures up to 450°F for 6 hours and 100 psi. With the assumption that surface scanning data can be obtained, an innovative rapid manufacturing technology solution is anticipated that will directly produce a female temporary tool for scarf repairs and demonstrate its ability to perform the following tasks: (1) lay-up and debulk of scarf repair structural plies with adequate vacuum integrity, (2) withstand cure temperatures for epoxy and/or bismaleimide carbon material systems (i.e. 350°F-450°F final temperatures) without experiencing splash tooling distortion, (3) withstand autoclave pressures up to 100 psi, (4) cure the structural repair patch via the provided process specification, and (5) ensure de-molded “hard” scarf repair patch meets inspection requirements (i.e. dimensional tolerances of ±0.005 inches). 

PHASE I: Develop a rapid manufacturing process that will directly produce a notional female temporary tool for hard patch fabrication. Demonstrate the tooling capability to perform tasks (1) through (5) listed above for an epoxy and/or bismaleimide scarf repair patch. Materials will not be provided by the Air Force in the Phase I. Identify the temporary tooling material’s durability and maximum upper use time at temperature. 

PHASE II: Further develop the mold tooling fabrication process along with the necessary surface scanning and computer aided design (CAD) models to rapidly manufacturing the temporary female form. Demonstrate the approach on representative complexity of a hard patch scarf repair for both epoxy and bismaleimide OML repairs as defined by the Air Force and OEM customers. Perform an assessment of the time/labor reduction associated with this process when compared to the multi-step male/female tooling designated in the standard structural repair process, such that mold fabrication time is reduced by 75%. 

PHASE III: Work with the program office to conduct all remaining technical tasks needed to implement an engineering change in standard Unit- and Depot-Level airframe structural solid-laminate repairs and commercialize the technology. 

REFERENCES: 

1: Baker, Alan. Development of a hard-patch approach for scarf repair of composite structure. No. DSTO-TR-1892. DEFENCE SCIENCE AND TECHNOLOGY ORGANISATION VICTORIA (AUSTRALIA) AIR VEHICLES DIV, 2006.

2:  Labor, J.D. and S.H. Myhre. Repair Guide for Large Area Composite Structure Repair. AFFDL-TR-79-3039. DEFENCE TECHNICAL INFORMATION CENTER, 1979.

KEYWORDS: Composite Scarf Repair, Additive Manufacturing, Composite Tooling, Structural Repair, Structural Adhesive Bonding 

CONTACT(S): 

Hilmar Koerner (AFRL/RXCC) 

(937) 255-9131 

hilmar.koerner.1@us.af.mil 

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