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Efficient On-Aircraft Composite Repair Process Requiring Minimal Support Equipment
Navy SBIR 2016.1 - Topic N161-017 NAVAIR - Ms. Donna Attick - [email protected] Opens: January 11, 2016 - Closes: February 17, 2016 N161-017 TITLE: Efficient On-Aircraft Composite Repair Process Requiring Minimal Support Equipment TECHNOLOGY AREA(S): Materials/Processes ACQUISITION PROGRAM: PMA 276 H-1 Helicopter Program Office OBJECTIVE: Develop an efficient on-aircraft repair process, requiring minimal support equipment, for structural components made of organic-matrix composite materials that can restore structural capabilities of those components. DESCRIPTION: Currently the Navy uses a Double Vacuum Debulk (DVD) process to repair damaged graphite/epoxy and fiberglass/epoxy composite components. This process works, however it only brings the component back to less than full strength. This process is done by impregnating dry fabric with resin, manually laying up the plies and debulking in a separate piece of support equipment. The debulked ply stack is then placed on the repair area and cured. This process, which limits the size of the repair, is expensive and creates logistical challenges when repairs are needed on aircraft in small detachments. A repair process is desired that provides repair patches of the same or better quality, but minimizes required support equipment and is less labor intensive. The proposed process should minimize the amount of support equipment required to reduce the logistics requirement. The repair patch should be a bonded organic-matrix composite that can be cured on aircraft, requiring no fasteners. The total time required to prepare and cure the repair patch should be minimized, and is not to exceed 8 hours. The repair process must be able to cure composite laminates at least 0.120 inches thick while minimizing porosity of the part, and the repaired laminate must be able to pass specified ultrasonic attenuation threshold requirements to ensure the porosity of the part is below 4%. The process must be able to produce patch sizes at least 15 by 15 inches, and the ability to produce larger patches is desired. The process must accommodate repairs on parts with complex curvature and variable shapes. The process must work on a radius of curvature as small as 4 inches or less, be able to be performed on parts oriented vertically or horizontally, and be able to be performed on sandwich panel configurations as well as laminated skins typical of airframe construction. The repair materials must provide properties equivalent to or better than the current wet layup DVD process using AS4 carbon fiber and EA 9390 resin system. The materials used in this repair should minimize low temperature storage requirements (i.e. typical freezer-type requirements) and maximize shelf life, with a desired shelf life of one year when stored at room temperature. At no point in the process can the structure being repaired be exposed to temperatures greater than 275 degrees F, and the final repaired product must have a wet glass transition temperature of at least 230 degrees F. The repair process must be able to be performed in an uncontrolled environment, such as in an aircraft hangar or outdoors. PHASE I: Design and determine the feasibility of a composite repair process that meets all of the topic criteria as discussed in the Description. Utilizing this proposed process and material, produce a flat panel that is a minimum of 15 by 15 inches and a minimum of 0.120 inches thick. This panel must be shown to pass an inspection that will verify the part has less than 4% porosity throughout. Provide initial substantiating data to show that the repair process can provide properties equivalent to or better than a DVD panel using AS4 carbon fiber and EA 9390 resin system. PHASE II: Develop and demonstrate the prototype composite repair process to ensure the process provides repeatable, reliable repairs of sufficient quality (4% porosity) and equivalent to or better mechanical/physical properties. Fabricate specimens for mechanical and physical testing using the developed process. Fabricate demonstration panels of the geometries and orientations outlined in the Description. The testing must include a limited set of screening tests sufficient to ensure acceptable properties and process repeatability that is to be developed by the company and approved by the government. Deliver any required support equipment. Provide a business case analysis of the process showing the savings that can be achieved as compared to the current DVD process. PHASE III DUAL USE APPLICATIONS: Complete mechanical coupon and element level testing sufficient to produce data to qualify the materials/process and determine material design allowable values for the use of the process. Complete an operational procedure to define all process requirements along with specifications detailing the requirements for all repair materials. Transition the repair process for Navy, fleet or commercial application use. This topic would apply to composite structures using a compatible material system that would need repair, including commercial aerospace applications along with any other commercial applications such as automotive or wind energy. REFERENCES: 1. T. H. Hou and B.J. Jensen, Evaluation of Double-Vacuum Bag Process for Composite Fabrication, Proceedings of 49th SAMPE International Symposium, 2004. 2. S. J. Ng, J. Brennan, E. Rosenzweig and T. Chen, Overview of Navy Repair Methods of Helicopter Composite Sandwich Structures, SAMPE, 2012. KEYWORDS: Composite; Composite Repair; Structural Repair; aircraft repair; carbon fiber; on-aircraft TPOC-1: 301-342-2181 TPOC-2: 301-342-2338 Questions may also be submitted through DoD SBIR/STTR SITIS website.
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