Air Force sbir 04. 1 Proposal Submission Instructions
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| PHASE I: Identify candidate binder and pigment materials to develop a high-temperature coating that meets the general specification requirements of the current camouflage coating (MIL-C-85285). Key coating performance parameters are thermal stability, color, gloss and compatibility with high-temperature structure substrates. The contractor shall identify candidate polymeric (nominal use temperature up to 700 °F) and preceramic (nominal use temperature up to 1200 °F) materials as candidate high-temperature binders. The contractor shall identify candidate pigments that will impart the proper optical performance. Trial coating shall be formulated, coated onto high-temperature structure substrates, and evaluated against the key performance parameters. Succession into Phase II will be based upon the results of this evaluation.
PHASE II: Further development of binder and pigment materials for improved thermal stability and coating performance. Processing techniques shall be developed with the assumption that individual parts will be coated, leading the way to innovative curing/heat-treatment techniques. Large-scale coating processes for entire aircraft are not to be pursued. A demonstration component of sufficient size (approximately 100 sq ft) will be coated at the end of this program, with limited thermal and coating performance testing on subelements. DUAL USE COMMERCIALIZATION: Similar high-temperature components on commercial aircraft. Industrial and automotive components may also benefit from these high-temperature coatings. REFERENCES: 1. Organic Coatings: Science and Technology, 2nd Ed., Eds. Wicks, Zeno, et al., ISBN 0-471-24507-0. KEYWORDS: high-temperature coating camouflage aircraft AF04-128 TITLE: Pilot Extraction Tool TECHNOLOGY AREAS: Air Platform OBJECTIVE: Develop universally applicable system that provides improved firefighter and rescue personnel access to cockpits and speeds pilot extraction under ground emergency conditions. DESCRIPTION: Advanced fighter/bomber aircraft require improved access for firefighting and rescue personnel to reach the cockpit. Emergency personnel must be able to open the canopy, accomplish engine shutdown, make the ejection system safe, and extract the pilot during ground emergencies. The proximity of engine inlets to the cockpit and the toughness of modern canopies prevents successful application of traditional approaches to pilot extraction. There is a significant risk of damage to high value-aircraft components and low-observable surface finishes during recurring live aircraft training involving simulated pilot extraction. The required system must be easily carried by one person, fire resistant, and compatible with aircraft and high-value aircraft parts. The system must also be safe for fire or rescue personnel to use for emergency situations (Note: Pilot extraction training and actual ground emergencies requires access to both sides of the cockpit.) PHASE I: Demonstrate the feasibility of the proposed tool via a prototype demonstration or subscale model demonstration and provide system concept for a potential Phase II effort. PHASE II: Build a full-scale system, delivered at the end of the Phase II effort to demonstrate its utility in tests with trained aircraft crash/rescue personnel on actual or closely simulated aircraft conducted by Air Force or contractor personnel. DUAL USE COMMERCIALIZATION: This tool would likely have commercial applications for emergency rescue personnel as well as DoD firefighting/rescue personnel. REFERENCES: 1. Air Force Document: Aerospace Emergency Rescue and Mishap Response Information (Emergency Services)TO 00-105E-9. Available at: http://www.robins.af.mil/ti/tilta/documents/to00-105E-9.htm KEYWORDS: fire, rescue, crash, open canopy, engine shutdown, safe ejection system, pilot extraction, ground emergency AF04-129 TITLE: Direct Manufacturing of Advanced Gas Turbine Engine Diffuser Cases TECHNOLOGY AREAS: Materials/Processes OBJECTIVE: Demonstrate a direct manufacturing technique for fabricating diffuser cases that significantly reduces the time and cost of development and production. DESCRIPTION: Nickel based alloys like INCO 718 and IMI 939 are widely used in the investment casting of diffuser cases for gas turbine engines. Modern engine designs have made these castings very complex and difficult to develop and produce at affordable costs with the first-time quality desired by OEM engine companies. Component manufacturing involves extensive, iterative development cycles resulting in undesirable levels of cost and time-to-first-article. The eventual product design is usually compromised to accommodate process difficulties and shortcomings such as shrinkage and inclusions associated with areas of abrupt section thickness changes and geometrical complexity. Therefore, an alternative method for fabricating diffuser cases is sought to reduce development cycle time and preclude these cost and quality issues. Several emerging direct manufacturing processes have the potential to greatly reduce first-article development cost and time and offer significant cost reductions in production. They also offer engineers the flexibility, during engine development, to easily and quickly modify designs to minimize weight while improving overall utility. However, these emerging direct manufacturing processes must be further developed for the manufacture of gas turbine components and rigorous economic analysis conducted to address both the technical and business questions associated with implementation and qualification to warrant industry support and investment. PHASE I: Demonstrate the engineering and economic feasibility of emerging direct manufacturing processes to fabricate diffuser cases from nickel based superalloys of interest to the gas turbine engine industry. The ultimate goal is to reduce first article development time and cost by over 50 percent, and to reduce final part production costs by 35 percent and product delivery cycle times by 50 percent. This phase will investigate the feasibility to directly manufacture sub-elements of typical diffuser case designs that represent anticipated areas of concern during fabrication. Subelement fabrications should include areas containing abrupt changes in cross-sectional thickness, sections containing hollow cavities, and elements containing bosses, flanges, or other types of attachments. Each element will be sectioned and characterized to evaluate dimensional and microstructural characteristics and for the presence of any metallurgical anomalies. Specific data will be gathered during the sub-element fabrication study to enable construction of an economic model of the process to project the savings offered and to provide clear understanding of the cost drivers and their respective magnitude of influence on overall process economics. An initial business case will be developed to understand the investment in equipment changeover and qualification expenses to warrant industry support and adoption. Additionally, a production capable supply chain scenario (from material suppliers to the delivery of a complete, machined component) will be proposed for further evaluation during the commercialization planning task in Phase II of the program. PHASE II: During Phase II of the program the process will be further developed to demonstrate full-scale fabrications of nickel based superalloy diffuser cases using production-capable direct manufacturing processes. Several first article cases will be fabricated to demonstrate process reproducibility under realistic production conditions. Cases will be heat treated, inspected using typical nondestructive evaluation and dimensional inspection techniques, and sectioned to document process capabilities and to establish an initial design data base. Portions of the fabricated cases will be further evaluated microstructurally to demonstrate acceptable material quality and metallurgical characteristics. General material property testing to obtain preliminary data per standardized testing techniques will be conducted to show the material acceptability for industrial use. Commercialization plans will be developed based on the supply chain scenario developed in Phase I. Qualification requirements will be established to introduce this new direct manufacturing process for fabrication of diffuser cases to the aerospace industry for production transition and qualification in Phase III. Deliverables shall include a detailed description of the processing path or paths used to create the prototypes plus one complete prototype component. DUAL USE COMMERCIALIZATION: Commercial applications shall include the manufacture of diffuser cases and other components for advanced gas turbine engines used for commercial aircraft and, potentially, for land-based turbines used in the power generation industry. REFERENCES: 1.D.L. Bourell and J.J. Beaman, Jr.,; "Freeform Fabrication - History and Current Processes," in Proceedings of Symposium on Rapid Prototyping of Materials, TMS Fall Meeting 2002, Columbus, OH, pp. 3 - 17, 2002. S.C. Danforth, D. Dimos, and F.B. Prinz (eds.), "Solid Freeform and Additive Fabrication - 2000," Proceedings of Materials Research Society Symposium, Vol. 625, 2000. J.W. Sears, "Solid Freeform Fabrication Technologies: Rapid Prototyping - Rapid Manufacturing," in International Journal of Powder Metallurgy, Vol. 37, No. 2, pp. 29 - 30, 2001. KEYWORDS: direct manufacturing, rapid manufacturing, solid freeform fabrication, diffuser cases, nickel-based superalloys AF04-130 TITLE: Ceramic Matrix Composites (CMCs) for Aircraft Brake Friction Materials TECHNOLOGY AREAS: Air Platform, Materials/Processes OBJECTIVE: Demonstrate and optimize CMCs as high-performance aircraft brake friction materials. DESCRIPTION: Carbon-carbon (C-C) composites are the state-of-the-art friction material for aircraft brakes. C-C composites suffer from a relatively low friction coefficient, a high variability in friction coefficient as a function of temperature, moisture content, pressure, susceptibility to oxidation at the use temperature, generation of a nuisance dust (through brake wear), and degradation from fluids commonly used on and around the aircraft. CMCs offer an alternative to C-C composites which may ameliorate many of these deficiencies, as well as provide for reduced wear rate and improved volumetric heat capacity. Silicon carbide (SiC)-based CMCs have been the primary focus of attention to date, and preliminary results have been promising [1-2], but other compositions may also be candidates. PHASE I: Optimize the CMC composition and brake configuration for maximum performance based on analysis and laboratory testing, and demonstrate the feasibility of the proposed brake material system through subscale testing. Assess the results, identify material and process refinements, and consider brake design modifications. PHASE II: Fabricate and dynamometer test at least two full scale military aircraft brake stacks over the course of the Phase II effort. The focus of the effort should be on 1) demonstrating desirable brake performance characteristics, 2) analytically evaluating brake design modifications to optimize brake performance with the CMC friction material, 3) improving/scaling up key fabrication steps, with an eye toward cost reduction and commercialization, and 4) Providing comparison of CMC versus C-C brake material characteristics (e.g. improvement in volumetric heat capacity of CMC versus C-C). DUAL USE COMMERCIALIZATION: The resulting advanced friction materials will be directly applicable to the large commercial aircraft brake market, where improved performance will result in reduced cost per landing. REFERENCES: 1. Vaidyaraman, Purdy, Walker, and Horst, "C/SiC Material Evaluation for Aircraft Brake Applications," in High Temperature Ceramic Matrix Composites, Eds. Krenkel, Naslain, and Schneider, pp. 802-808, Wiley-VCH, Weinheim, Germany (2001). 2. Heidenreich, Renz, and Krenkel, "Short Fibre Reinforced CMC Materials in High Performance Brakes," in High Temperature Ceramic Matrix Composites, Eds. Krenkel, Naslain, and Schneider, pp. 809-815, Wiley-VCH, Weinheim, Germany (2001). KEYWORDS: aircraft Brakes, ceramic matrix composites (CMCs), friction materials, C/SiC (carbon fiber reinforced silicon carbide), process development AF04-131 TITLE: Erosion Protection Materials for High-Temperature Composites TECHNOLOGY AREAS: Air Platform, Materials/Processes OBJECTIVE: The objective of this research effort is to develop and demonstrate new erosion protection material systems for use on high-temperature polymer matrix composite (HTPMC) materials in aircraft turbine engines. DESCRIPTION: HTPMCs have tremendous potential to reduce the weight and cost of military aircraft turbine engines if producibility and durability in service environments can be demonstrated. An important factor for engine durability is resistance to ingested particle erosion. This requires that viable erosion protection materials can be developed. Erosion protection coating materials have not been demonstrated that are suitable for use in hot engine components with service temperatures of up 650 °F for the life of the engine. This research effort is to develop affordable coating systems (materials and processes) that are compatible with candidate HTPMCs such as Avimid N, PMR-II-50, AFR-700, and AFRPE. PHASE I: Candidate materials and processes will need to be screened for: 1) erosion protection at engine service temperatures, and 2) compatibility with HTPMC materials and processes used for turbine engine components. For the Phase I effort coupon-level evaluation is appropriate. Erosion protection requirements should be coordinated with engine manufacturers for specific evaluation methods and pass/fail criteria. Selection of specific HTPMC substrate materials should be coordinated with engine manufacturers to ensure the relevance and suitability of erosion protection materials developed under this proposed research effort. This Phase I screening should result in down-selection of best performing erosion protection concepts that will be further developed under a Phase II effort. A feasibility study shall be performed as part of Phase I evaluating implementation issues including manufacturing and material development costs. PHASE II: The selected materials and processes will be fully optimized and scaled up to demonstrate engine service environments. These environments include durability, affordability and erosion resistance. All critical engine environmental durability factors should be addressed and demonstrated. These requirements must be determined through coordination with engine manufacturers. DUAL USE COMMERCIALIZATION: Commercial aircraft, rotorcraft, and power generation turbines have similar durability and temperature performance requirements to military turbine engines. These commercial applications will benefit from coating technology that will enhance erosion protection of composite components. REFERENCES: 1. "Mechanical Testing of PMCs Under Simulated Rapid Heat-Up Propulsion Environments (1. Temperature Measurement)", SAMPE Proceedings, May 12-16, 2002, Long Beach, CA KEYWORDS: erosion protection, polymer matrix composites, turbine engines, coatings AF04-132 TITLE: Encapsulated Resin for Non-Autoclave Resin Film Infusion Composites Repair TECHNOLOGY AREAS: Air Platform, Materials/Processes OBJECTIVE: Develop processing techniques for on-aircraft repair of composite components using encapsulated resins that can be stored at room temperature. DESCRIPTION: Repair of composite aircraft structure in the field generally requires an adhesive bonding approach to provide the load transfer and restore the original design strength of the composite laminate. These materials generally require freezer storage and have limited shelf life. Heat and pressure are required to cure the adhesive and patch materials and obtain a uniform, nonporous adhesive layer. Resin film infusion (RFI) is a process in which calendared film or plaques of resin are placed on top of or below fiber preforms and then consolidated and cured together under vacuum-only conditions. Achieving the high performance characteristics has not been realized as yet by RFI due, in large part, to difficulty in achieving void-free parts. Air entrapped in the resin films cannot be removed from the resin film when in its solid form, and once the resin melts, poor flow characteristics lead to an inability to evolve the air from the consolidating laminate. Encapsulating a predetermined amount of resin and applying vacuum prior to flow offers a solution to the limitations of the RFI process. Innovative approaches will improve the RFI process and reduce costs for aircraft repair. The program goal is to develop a method to use materials that can be stored at room temperature and which bond with existing adhesive materials (for a 250 to 275°F bond temperature), although other materials and processes may be considered. With the advent of the Air Expeditionary Force concept within the Air Force, increased emphasis is being placed on quick reaction forces requiring minimal support equipment and materials not having short shelf lives nor requiring environmentally controlled storage conditions. PHASE I: Identify candidate materials and processes and demonstrate resin film infusion processing procedures to achieve mechanical and physical properties with a decrease in repair cure time. The contractor shall demonstrate, at a laboratory scale, properties that are comparable to existing autoclave cured (350F) epoxy materials and withstand a minimum service temperature of 180F. PHASE II: Refine and optimize the process investigated during Phase I. The repair process shall be demonstrated and strength tested on a US Air Force representative composite structure with equipment and skill level compatible to a field base location. The term process, when applied to the repair process, is meant to include the basic repair procedures, any support equipment (i.e., heating equipment, vacuum bagging materials, etc.), and the environmental considerations they entail. DUAL USE COMMERCIALIZATION: The Air Force has a variety of aircraft applications that a successfully developed material would find use in. Commercial applications include recreational sporting industry. REFERENCES: 1. MIL-HDBK-337, 1982-12-01, Adhesive Bonded Aerospace Structure Repair 2. MIL-HDBK-17/1F, Composite Materials Handbook Volume 1, Polymer Matrix Composites Guidelines for Characterization of Structural Materials 3. MIL-HDBK-17B(1), Polymer Matrix Composites, Volume 1. Guidelines (S/S by MIL-HDBK 17/1) 4. MIL-HDBK-17/2F, Composite Materials Handbook, Volume 2, Polymer Matrix Composites Materials Properties KEYWORDS: composites, repair, processing, nonautoclave AF04-133 TITLE: Damage Detection in Composites via Passive Monitoring Techniques TECHNOLOGY AREAS: Materials/Processes OBJECTIVE: Develop a graphite fiber composite damage detection method which improves rapidity, cost, and/or certainty of post-damage inspections at different levels of depth. DESCRIPTION: Damage in graphite fiber composites can be detected in numerous ways. The objective of this program is to detect damage easily. The techniques proposed may be for field and/or depot maintenance. They should not involve real-time acquisition schemes, but rather, schemes that embrace either presently unused detection techniques, or schemes that make present methods faster and/or more efficient. Examples that will be strongly considered include, but are not limited to: 1) Pressure-sensitive paint. Paint responds to impact by changing color due to microencapsulants; possibly different impact levels might trigger different colors. Color change readily indicates impact events to field and/or Air Logistic Center personnel, who can then focus attention on detailed inspection of affected areas. Paint must be compatible with the usual requirements on competing aircraft paint systems, as well as permit maintenance crews to work on the aircraft without triggering the paint impact detection properties. 2) Rapid application of ultrasound coupling medium. Most of the time involved in ultrasonic inspection is associated with assuring the presence of water (usual coupling medium) between the body to be inspected and the ultrasonic transducer. Methods exist, and/or can be improved, which dramatically reduce the time associated with insuring proper application of couplant (normally water). 3) Proposed methods involving electrical changes [5] due to graphite fiber or matrix strains, or eddy current phenomena, will be considered only if accompanied by very convincing evidence of suitability to full-scale aircraft application by the end of Phase II. Such evidence will necessarily be based on the maturity of such technology at the time of the Phase I proposal submission. 4) Radio-opaque embedded fiber strategies are not encouraged due to concerns about potential needs to re-qualify materials systems, which in general would be cost prohibitive. PHASE I: Demonstrate damage sensing (and possibly strain sensing) capabilities at lab environment level, for systems of interest to the Air Force such as IM7/5250-4 and IM7/977-3 in laminated tape and woven forms. Deliverables include all data generated, and reports regarding progress on the concept and feasibility of meeting the Phase II objectives. Early and continuing interactions with sponsors are recommended. PHASE II: Develop and deliver a technology approach which can produce a product that can be inserted in an Air Force system program with a minimum of additional effort, preferably within 2 years following Phase II completion. A production-capable system for application to aircraft is thus desired. A prototype will be a required deliverable, in addition to the usual data and reports. Collaboration in this phase with aircraft manufactureres and/or Air Force personnel with systems access is highly encouraged. DUAL USE COMMERCIALIZATION: This technology can be applied to damage sensing in graphite fiber-reinforced structures, e.g., in civil aircraft, armor, advanced vehicle powertrains, off-shore drilling platforms, Naval or civilian marinecraft, etc. Collaboration in this phase with aircraft manufacturers and/or Air Force personnel with systems access is highly encouraged. REFERENCES: 1. Light, G.M., 1986, Development of Encapsulated Dye for Surface Impact Damage Indicator System, Phase II: Southwest Research Insititute SwRI Project No. 17-1056. More info at http://www.swri.edu/4org/d14/ndedept/home.htm H. W. Schlameus, G. L. Light, and C. H. Parr, Impact Indicator Coatings, Second DOD/NASA Composites Repair Technology Workshop, San Diego, CA, November, 1986. More info at http://www.swri.edu/4org/d01/appchem/encap/home.htm and http://www.swri.edu/3pubs/papers/d01/01pres.htm Crane, R., 1981, Composite Structures and Method of Detecting of Mechanical Damage Thereto: Assigned to USAF, USA, Patent No. 4,255,478. Buynak, C.F. and Crane, R.L., 1987, A Novel Acoustic Coupling Device using Permeable Membrane: Materials Evaluation, Vol. 45, No. 6, pp. 743-746. KEYWORDS: damage, composite, graphite, paint, ultrasound AF04-134 TITLE: Nondestructive Inspection (NDI) of Fastener Holes in Thick Multi-Layer Structure TECHNOLOGY AREAS: Materials/Processes OBJECTIVE: Develop NDI to detect cracking within Taper Lok fastener holes of thick multilayered structures without the removal of fasteners. DESCRIPTION: A requirement exists to conduct NDIs on a thick multilayer wing structure. Unfortunately, due to the thickness and complexity of the structure, there is currently no means of inspecting this hardware without removing the Taper Lok fasteners and conducting bolthole eddy current inspection. Removal of the Taper Lok fasteners would cause damage to the fastener bore and result in cost prohibitive bore repair and fastener reinstallation. 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