Air force 14. 1 Small Business Innovation Research (sbir) Proposal Submission Instructions
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PHASE II: Design and fabricate prototype-scale equipment for long-length conductor manufacturing, and provide deliverables of conductors 1-10 meters long rated for current in the range <=20 Amps. Evaluate properties such as mechanical, uniformity as a function of length, fatigue and time-stability of such conductors, and determine whether they can meet military specifications. Provide a cost benefit analysis for a specific AF system of interest such as BAO Kit or SUAS. Present results at WPAFB. PHASE III DUAL USE APPLICATIONS: A multitude of aerospace vehicles will benefit from reliable light-weight conductors that will improve fuel efficiency and increase payload capability. Other potential civilian users include law enforcement, rescue crews, and others who typically have to carry electronic equipment to do their jobs. REFERENCES: 1. M. Meyyappan, “Carbon Nanotubes: Science and Applications,” CRC Press LLC: Boca Raton, FL, 2005. 2. Lan,Y., Wang,Y., and Ren, Z.F., "Physics and applications of aligned carbon nanotubes," Advances in Physics 60, 553-678 (2011). 3. Y. Zhao, et al., “Iodine doped carbon nanotube cables exceeding specific electrical conductivity of metals,” Scientific Reports / 1:83 / DOI: 10.1038 / srep00083. 4. I. Khrapach, et al, “Novel highly conductive and transparent graphene based conductors,” arXiV: 1206.0001v1. 5. M. Ezawa, “Dirac theory and topological phases of silicon nanotube,” http://arxiv.org/abs/1203.4654v1. KEYWORDS: mobile electronics, wearable electronics, battlefield air operations, lightweight electrical conductor, mass specific electrical conductivity, carbon nanotube, silicon nanotube, carbon fiber composites, metal composite, graphene, graphene multilayer composite, highly oriented pyrolytic graphite, topological insulator, electrical power systems, power transmission wires and cables AF141-063 TITLE: Modeling the Impact of Silica Particle Ingestion on Turbomachinery Life KEY TECHNOLOGY AREA(S): Materials / Processes The technology within this topic is restricted under the International Traffic in Arms Regulation (ITAR), 22 CFR Parts 120-130, which controls the export and import of defense-related material and services, including export of sensitive technical data, or the Export Administration Regulation (EAR), 15 CFR Parts 730-774, which controls dual use items. Offerors must disclose any proposed use of foreign nationals (FNs), their country(ies) of origin, the type of visa or work permit possessed, and the statement of work (SOW) tasks intended for accomplishment by the FN(s) in accordance with section 5.4.c.(8) of the solicitation and within the AF Component-specific instructions. Offerors are advised foreign nationals proposed to perform on this topic may be restricted due to the technical data under US Export Control Laws. Please direct questions to the AF SBIR/STTR Contracting Officer, Ms. Kristina Croake, kristina.croake@us.af.mil. OBJECTIVE: Develop a decision support tool to determine the impact of silica particle ingestion on component service life of the engine hot section. DESCRIPTION: Operational requirements of commercial and military aircraft often render the traditional method of total avoidance of silica-rich particle contaminate ingestion infeasible. The 2010 Iceland volcanic ash cloud was of such a large extent that military sorties had to be rerouted around the outer perimeters of the cloud. Many airports around the world are subject to long-term exposure to sand-laden air. Also, volcanic clouds containing silica can be encountered in-flight before detection, resulting in inadvertent engine exposure. The current total avoidance strategy is driven by lack of detailed knowledge of how the two common particle types, fine sand and volcanic ash, affect engine components through several mechanisms of damage. These include erosive wear from mechanical action, impact damage from hard particles, contamination of accessories systems, and accumulation from melting and subsequent deposition within the hotter gas turbine engine components. Catastrophic failure of turbomachinery from high levels (approximately 0.1 gm dust per cubic meter of air) of calcium-magnesium-aluminum-silicon oxide (CMAS) particulates has been documented, but the effect of shortened component life imposed by reduced particulate loads that allow continued engine operation (approximately 1 µg to 1 mg per cubic meter), is not well understood. Glass accumulation on the combustor and turbine components may be an event that does not mandate replacement if thermal cycling allows thermal spallation of the glass accumulated; however, this may compromise component life. Compounding the problem is the widespread use of thermal barrier coatings in the turbine and combustor sections of advanced engines, where CMAS-induced coating loss can increase the risk of thermal fatigue in the base materials. Since sand and volcanic ash composition vary widely with geographic location, a tailorable CMAS particulate may be worth deriving for experimental approaches. The various minor chemical components in the CMAS can alter the phase change temperature of the silica, as well as serve as fluxing agents. It will be necessary to determine the depositional locations as well as the combined effects of some of the anticipated component life-reduction mechanisms, including degradation of the thermal barrier coating (TBC) layers, blade harmonic loading, and cooling boundary layers. This topic seeks a novel method of accounting for the cumulative exposure effects of silica-particulate containments upon turbomachinery components in regards to time change interval reduction from the phase-change deposition of the particulates. An expert-system and/or decision support tool that allows for reduction of operable time for various stages based on particulate size, concentration, and exposure time is a desired deliverable product. Phase I shall focus on the experimental design concept and numerical modeling tools that will use the experimental results to build a service life estimation tool that will downgrade the remaining service life based on CMAS accumulation, cooling disruption, thermal barrier erosion and other compromised conditions. Phase II will involve testing components that have had CMAS accumulated on them, validating the numerical extrapolation model against them to generalize the change in loading as a function of CMAS time and concentration exposure. Based on the outcomes of this test effort, a validated fatigue-failure numerical model will be used to determine remaining service life as a function of CMAS exposure levels. PHASE I: Demonstrate the feasibility of simulating glass accumulation effects on components at various conditions within gas turbines. Demonstrate CMAS formulations for representative sand and volcanic ash systems. Show a modeling concept that can be validated against test conditions and allows a generalized component time change modification model to be generated. Develop a test plan for this purpose. PHASE II: Demonstrate testing for glass accumulation effects of CMAS particulates at various engine operating conditions (temperature, species, velocity, concentration, and size parameters). Validate the modeling concept against test data from demonstration cases in collaboration with government and industry end users. Deliver a decision support tool that adjusts time change interval of components based on silica particulate exposure history and engine conditions, and an instruction manual for tool use. PHASE III DUAL USE APPLICATIONS: Adoption of the tool by aircraft maintenance organizations will support dispositions for exposed engines. Commercial benefits include improved opportunities for providers of air services to refine risk when operating in environments with pervasive silica-particulate contamination. REFERENCES: 1. Guffanti, M., Casadevall, T., and Budding, K., “Encounters of Aircraft with Volcanic Ash Clouds: A Compilation of Known Incidents, 1953–2009,” USGS Data Series 545, Reston VA, 2010. 2. Grindell, T. and Burcham, F., “Engine Damage to a NASA DC-8-72 Airplane From a High-Altitude Encounter With a Diffuse Volcanic Ash Cloud,” NASA Glenn Research Center, NASA/TM-2003-212030, August 2003. 3. Casadevall, T. and Murray, T., “Advances in Volcanic Ash Avoidance and Recovery,” Boeing Commercial Airlines Group. 4. Manual on Volcanic Ash, Radioactive Material and Toxic Chemical Clouds, Doc 9691 AN/954, International Civil Aviation Organization, Second Edition, 2007. 5. Hamed, A., Tabakoff, W., and Wenglaz, R., “Erosion and Deposition in Turbomachinery,” Journal of Propulsion and Power, Vol. 22, No. 2, March–April 2006. 6. Mechnich, Peter; Braue, Wolfgang; and Schulz, Uwe. " High-Temperature Corrosion of EB-PVD Yttria Partially Stabilized Zirconia Thermal Barrier Coatings with an Artificial Volcanic Ash Overlay", Journal of the American Ceramic Society, 94 [3] 925-931 (2011). (uploaded in SITIS 12/20/13) 7. British Airway Flight 9 Ash Encounter, Wikipedia, Aug. 1, 2013. (uploaded in SITIS 12/20/13) KEYWORDS: decision support, service life modification, trade space, particulate contamination, gas turbine, maintenance disposition, volcanic ash, sand ingestion, glass AF141-064 TITLE: Additive Metal Manufacturing (AMM) Process Development for Gas Turbine Engine Component Repair
OBJECTIVE: Develop and validate an additive-metal manufacturing repair process for complex engine components in aging Air Force (AF) fleets. DESCRIPTION: Gas turbine engine components experience damages such as fatigue, foreign object damage (FOD), erosion, and fretting wear that make the sustainment of fleets burdensome. Damaged components in most AF fleets are usually replaced with new parts supported by the original equipment manufacturer (OEM); however, the AF still owns and operates a number of aging engines no longer supported by OEMs. Though some of the unsupported components were manufactured from known, commercially available materials, some were made from either proprietary or unavailable material compositions with sometimes unknown manufacturing processes and parameters. In order to sustain aging fleets and unsupported components, an additive metal manufacturing (AMM) and re-engineering effort for part repair and remanufacturing is being explored. The primary manufacturing procedure for this effort is direct metal laser sintering (DMLS). This procedure requires both test validation and equipment enhancement to qualify the AMM repair and remanufacturing capability for flight components. Validating the remanufacturing or repair of unsupported engine parts requires adherence to the Federal Aviation Administration (FAA) part manufacturing approval (PMA) process. The FAA states that two basic ways that a PMA applicant can show that a part meets airworthiness standards are as follows: first, establish identicality between the remanufactured or repaired and certified parts, and second, show through tests and computations that the remanufactured or repaired part meets airworthiness requirements. In the case of identicality, it must be shown that the remanufactured or repaired part is identical in material composition and functionality to the original part via mechanical material property performance. As for test and computation, the remanufactured or repaired part must show similarities in airworthiness functionality to the original part, only; meaning, the identicality of the functionality and the material composition is not of great importance. The innovative research required to attain airworthiness and to meet the standard for AMM repair is as follows: for Phase I, develop a DMLS repair process for blended/cutout aluminum and titanium coupon specimens and validate functionality by comparing failure results of repaired and baseline (pristine) coupon specimens; for Phase II, validate the DMLS repair process for an aluminum alloy with the same functionality as the TF33 fuel pump housing and conduct the repair on the part. Material property data for TF33 will be provided by the AF. Phase II may also require an enhancement to the equipment used for DMLS repair, specifically software control and powder application, so that repairs can be made on components with geometries that are more complex than that of coupon specimens. A final presentation of the repair capability demonstrated in the Phase II activity is required at Wright-Patterson Air Force Base. To successfully perform the work described in this topic area, offerors may request to utilize unique facilities/equipment in the possession of the U.S. Government located onsite at Wright-Patterson Air Force Base. Accordingly, the following items of Base Support may be provided to the successful offeror, subject to availability and negotiations, in accordance with the clause in Air Force Materiel Command FAR Supplement (AFMCFARS) 5352.245-9004 “Base Support.” The facilities/equipment include the Turbine Engine Fatigue Facility (TEFF) and certain fracture, fatigue, and vibration testing and measurement systems therein. PHASE I: Validate high and low strain rate material strengths [1,2], fatigue crack growth [3], low cycle and high cycle fatigue rates [4,5] of DMLS-repaired aluminum and titanium coupon specimens via comparative experimental study against respective baseline specimens. Demonstrate repair capability on coupon specimens by showing functionality within 20 percent when compared to baseline results. PHASE II: Determine material properties of the TF33 fuel pump housing and gear wipe (AFRL), conduct Phase I activities on housing and gear wipe materials, and conduct successful repair of housing and gear wipe. Deliverables: a FAA PMA validated repair capability for TF33 housing and gear wipe. PHASE III DUAL USE APPLICATIONS: Military: AMM repair used to reduce sustainment costs, including part purchasing, and the design and fabrication of obsolete parts. Commercial: AMM repair is adopted by OEMs or a nonaerospace consumer--the former for solidifying customer satisfaction, and the latter to reduce part replacement costs. REFERENCES: 1. American Society for Testing and Materials, “E 8 - 09: Standard Test Methods for Tension Testing of Metallic Materials,” ASTM Book of Standards, 2009; Vol. 03.01, ASTM International, West Conshohocken, PA. 2. T. Nicholas, "Tensile Testing of Materials at High Rates of Strain," Exp. Mech. 21 (1981) pp. 177-188. 3. American Society for Testing and Materials, “E 647 - 08: Standard Test Method for Measurement of Fatigue Crack Growth Rates,” ASTM Book of Standards, 2008; Vol. 03.01, ASTM International, West Conshohocken, PA. 4. American Society for Testing and Materials, “E 466 - 07: Standard Practice for Conducting Force Controlled Constant Amplitude Axial Fatigue Tests of Metallic Materials,” ASTM Book of Standards, 2009; Vol. 03.01, ASTM International, West Conshohocken, PA. 5. George, T., Seidt, J., Shen, M.–H.H., Cross, C., and Nicholas, T., “Development of a Novel Vibration-Based Fatigue Testing Methodology,” Int. J. of Fat., 2004, Vol. 26, pp. 477-486. KEYWORDS: sustainment, repair, additive metal manufacturing, DMLS, aluminum, TF33 fuel pump AF141-065 TITLE: Structural Health Monitoring (SHM) Methods for Aircraft Structural Integrity KEY TECHNOLOGY AREA(S): Air Platforms OBJECTIVE: Develop fuse-like SHM techniques for the ASIP environment. Methods must be reliable, low cost, and durable. Methods must reduce maintenance burden, while maintaining safety. DESCRIPTION: The U.S. Air Force utilizes a damage-tolerant design approach to ensure the structural safety and reliability of the airframes on its fleet. A critical facet of this damage-tolerant design approach is nondestructive inspection (NDI). These inspections ensure that a critical crack is not present in the inspected region and that any defects present are so sufficiently small that they will not grow to failure during the next service interval. While this approach is effective, the maintenance burden associated with repeated inspections can be problematic. The inspections themselves can be difficult and significant disassembly of the structure may be required to gain access to the inspection areas. Since there are typically many inspection points on an aircraft, this can result in significant cost and downtime for the aircraft. Unconventional approaches to supplement these aircraft structural inspections in an ASIP environment are desired. This effort will focus on developing a fuse-like system that will be implemented in the area where cracking is anticipated to occur on a metallic structure. These mechanical fuses may be bonded or utilize a direct-write methodology in which the sensor traces are directly deposited in the surface of the structure. Once the crack initiates and starts to propagate through the region where the sensors have been attached, the sensor traces will break, thus giving the indication of damage. The developed fuse system should be able to detect cracking of 0.1 inch or better. Remote determination functionality must permit indications of initiation and size of damage, either through a strategically located data port or wirelessly transmitted to a hand-held device. The developed sensor technology must be suitable for robust operation in the austere flight/field environments with limited maintenance and low false call rates as well as high detection capability. In Phase I, demonstrate the developed fuse-type SHM approaches for aircraft structural integrity to provide a proof-of-concept model or prototype that is scalable to a full-size aerospace structure. Crack initiation and sizing will be demonstrated on a simple metallic, laboratory dogbone specimen. Demonstrate that the approach causes no degradation in the underlying specimen. Provide the ability to perform sensor health checks to differentiate between a damaged sensor or a damaged structure. Outline a viable validation path to ensure compliance with MIL-HDBK-1530C (ASIP) or a logical evolution thereof. In Phase II, develop and conclusively demonstrate a prototype application of the developed approaches and validate the reliability of such a device in a relevant environment. It is recommended to partner with a potential end user of the technology to maximize the relevance of the demonstration and facilitate subsequent transition/commercialization. The demonstration of the developed system should show that the system is applicable to a more realistic aircraft structure with significantly more complicated geometry. Demonstrate that the developed technique can achieve a high probability of detection rate and low probability of false alarms. Examine the possibility of moving from accessing the data via a strategically located data port to the use of a hand-held device where the data could be wirelessly transmitted. PHASE I: Demonstrate fuse-type SHM methods for aircraft structure to prove scalability. Crack detection and sizing will be demonstrated on metallic laboratory specimens. Demonstrate that the approach causes no negative impacts. Sensors must differentiate between damaged sensors and damaged structure. Outline a viable path to ensure compliance with MIL-HDBK-1530C. PHASE II: Develop and demonstrate a prototype application and demonstrate the reliability of the device in a relevant aircraft environment. Recommend partnering with a potential end user to maximize the relevance and facilitate subsequent transition. The demonstration should show that the system is applicable to realistic aircraft structures with complicated geometry. Demonstrate high probability of detection (0.1 inch or better) and low probability of false alarms. PHASE III DUAL USE APPLICATIONS: The developed fuse-type SHM system would be fielded for relevant U.S. Air Force applications and fielded into broader commercial markets. REFERENCES: 1. MIL-HDBK-1530C, General Guidelines for Aircraft Structural Integrity Program (ASIP). 2. Air Vehicle Integration and Technology Research (AVIATR) Delivery Order 0002: "Condition Based Maintenance Plus Structural Integrity (CBM + SI) Strategy Development," Final Report, Nov 2010, DTIC Number ADA546937. 3. Air Force Institute of Technology, Thesis for Master of Science in Systems Engineering, "An Enhanced Fuselage Ultrasound Inspection Approach for Integrated Structural Health Monitoring Purposes," Mar 2012. KEYWORDS: health monitoring, structural integrity, nondestructive inspection, damage tolerant design, lifecycle management, aircraft availability, total ownership cost AF141-066 TITLE: Use more accurate aircraft usage data in predicting life and scheduling inspections KEY TECHNOLOGY AREA(S): Air Platforms OBJECTIVE: Obtain a more accurate prediction of remaining life and inspection interval for an individual aircraft by converting actual aircraft usage data into stresses on the structure via physics-based, real-time aeroservoelastic simulations. DESCRIPTION: The process of determining initial or remaining aircraft structure life has not significantly changed in 50 years. It is still a highly manual and labor intensive process, individual steps are not easily integrated together, and the advantages of high performance computing have not been fully utilized. Recently, the Air Force Research Laboratory has produced a long-term vision, called Airframe Digital Twin (Ref 1), that is beginning to address these issues. This project will be one of the crucial early steps toward the Airframe Digital Twin vision.
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