Room Temperature Shelf-Life Pre-Impregnated Carbon Fiber Fabric for use in Out-of-Autoclave Aircraft Repair
Navy SBIR 2019.2 - Topic N192-084
NAVAIR - Ms. Donna Attick - email@example.com
Opens: May 31, 2019 - Closes: July 1, 2019 (8:00 PM ET)
TECHNOLOGY AREA(S): Materials/Processes
ACQUISITION PROGRAM: PMA261 H-53 Heavy Lift Helicopters
OBJECTIVE: Develop an out-of-autoclave processable, pre-impregnated carbon fiber fabric that has a room temperature shelf life, is curable at low temperatures, and performs equal to or better than the materials currently being used for repair on Navy platforms.
DESCRIPTION: The resin pre-impregnated fabrics (pre-pregs) the U.S. Navy currently uses require storage at or below freezing. This requirement drives up sustainment cost and limits the ability to perform certain types of Organizational level (O-level) repairs where freezer storage is not readily available. The fabrics also must be cured in an autoclave or through a Double Vacuum Debulk (DVD) procedure, which drives the need for expensive equipment to support repairs and also limits the location of where repairs can be performed. Only a few commercially available room temperature storage pre-pregs can be cured outside of an autoclave but these materials need to be cured at relatively high temperatures (>250°F) and frequently yield high porosity laminates. The required processing exposes the parent materials to conditions outside their operational temperature windows, which can result in degradation of material properties. Additionally, higher porosity causes poor laminate quality and can result in premature part failure.
The desire is to produce a pre-preg that reduces the use of cold storage and equipment needed for cure, while producing a laminate of sufficient quality. The pre-preg would be expected to meet the following requirements:
- Can be produced as a plain woven and an unidirectional carbon fiber fabric
- Minimum shelf life of 1 year when stored in a hangar (100°F), but longer is preferred
- Reasonably tacky in order to perform repairs on part surfaces oriented vertically or horizontally
- Reasonably drape-able to form over complex curvatures with as small as a 4 inch radius or less
- Able to achieve a cure percentage of at least 95% when cured on aircraft
- Can be cured in an uncontrolled environment, ideally but not limited to 45-65% humidity at 65-75°F.
- Minimize the use of equipment needed to cure
- Cure time of 2.5 hours or less
- Cure cannot expose the part to temperatures greater than 200°F although as low as 150°F would be preferred.
- Porosity of laminate less than 4% by volume
- Wet glass transition temperature (Tg) of at least 230°F, but a higher wet Tg is desirable
- Exposure to common aircraft fluids should not cause degradation of mechanical properties greater than 11% of the original strength. Common aircraft fluids include, but are not limited to anti-icing fluid, runway deicers, electronic equipment coolant, hydraulic fluid, lubricating oil, jet fuel, turbine fuel, aircraft cleaner, MEK, and acetone.
-Must be capable of being co-cured and bond with another epoxy-based adhesive system.
- Ability to procure material in small quantities (by the roll) is desirable
Threshold Composite Laminate Mechanical Properties
- 0° tensile strength of 114 ksi (Room Temp), 109 ksi (180°F Wet)
- 0° compression strength of 69 ksi (Room Temp), 48 ksi (180°F Wet)
- 0° short beam shear strength of 8.9 ksi (Room Temp), 5.7 ksi (180°F Wet)
Objective Composite Laminate Mechanical Properties
- 0° tensile strength of 158 ksi (Room Temp), 151 ksi (180°F Wet)
- 0° compression strength of 130 ksi (Room Temp), 97.1 ksi (180°F Wet)
- 0° short beam shear strength of 12.7 ksi (Room Temp), 8.7 ksi (180°F Wet)
- OHT (open hole tension) strength of 57 ksi (Room Temp), 56 (180°F Wet)
- OHC (open hole compression) strength of 52 ksi (Room Temp), 43 (180°F Wet)
- CAI (compression after impact) strength of 44 ksi (Room Temp)
PHASE I: Design and determine the feasibility of developing a pre-preg as outlined in the Description. Design a proposed resin system and determine the feasibility of the resin system meeting the Tg requirements. Show feasibility of meeting the shelf life requirements as outlined in the Description. The Phase I effort will include prototype plans to be developed under Phase II.
PHASE II: Develop and provide a prototype pre-preg and demonstrate that it will produce a laminate of sufficient quality as outlined in the Description. Produce a resin system and fabricate a pre-preg with the resin system.
Fabricate specimens for mechanical and physical testing using the developed pre-preg. Conduct, in coordination with the Government, testing that includes a limited set of screening tests sufficient to ensure acceptable properties.
PHASE III DUAL USE APPLICATIONS: Transition technology to platforms/industry after verifying the material meets program specific requirements and all the performance requirements as outlined in the Description. The private aerospace sector, along with any small composite fabrication shops, will also have interest in this technology not only for repair but for primary structures. Room temperature shelf life would eliminate the need for freezer storage thus reducing the logistical footprint. It would also significantly extend the working life of the material, which would allow for the fabrication of larger parts without pushing the materials out time envelope. A capable, out of autoclave material would reduce the cost associated with composites fabrication by eliminating expensive autoclave operation. Materials could be cured using a conventional oven which would open composite fabrication to more companies. If the material is developed to reach the processing and mechanical properties in the Description section, it would be applicable to a wide variety of aircraft and repair types. This would bring down support costs for both military and the commercial aircraft sector, allowing autoclave quality repairs to be done closer to their fleet.
1. Guard, C., Hamnett, M., Neumann, C., Lander, M., and Siegrist, H. “Typhoon Vulnerability Study for Guam.” Water and Environmental Research Institute of the Western Pacific: Guam, 1999. http://www.weriguam.org/docs/reports/85_1.pdf
2. Hamill, L., Centea, T., Nilakantan, G., and Nutt, S. “Surface Porosity in Out-of-Autoclave Prepreg Processing: Causes and Reduction Strategies.” SAMPE Tech.: Seattle, 2014. https://www.researchgate.net/publication/267333388_Surface_Porosity_in_Out-of- Autoclave_Prepreg_Processing_Causes_and_Reduction_Strategies
KEYWORDS: Room Temperature; Pre-impregnated; Out of Autoclave; Low Temperature Cure; Repair; Composites; Organizational level repairs; Double Vacuum Debulk