Air force 12. 1 Small Business Innovation Research (sbir) Proposal Submission Instructions
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| The material should provide sufficient EM attenuation to produce internal field strengths 20 dB (threshold) /32 dB (objective) below the typical electronics immunity levels and ionizing radiation fluence below circuit upset levels. These levels are defined by MIL-STD-464C and MIL-STD-2169B.
The appendix of MIL-STD-2169B is classified and only available to personnel with a Secret level clearance. Requirements for the Phase II are defined by MIL-STD-2169B and its appendix. Only contractors able to obtain a Secret level security clearance should submit proposals against this topic. PHASE I: Establish feasibility and technical merit of proposed solution through modeling, design analysis, and laboratory experiments on coupon-level samples specifically focusing on performance against the MIL-STD 464 waveform. Assess any cost, performance, or manufacturability issues and recommend risk reduction activities to address them. PHASE II: Manufacture and test a prototype enclosure/test article defined by the US Air Force sponsor to verify manufacturability and performance against the classified MIL-STD-2169B waveform. The hardened composite technology Technical Readiness Level (TRL) should be at 6 to 7 and the Manufacturing Readiness Level (MRL) should be at 3 to 4 by the end of Phase II. Only contractors able to obtain a Secret level security clearance should apply. PHASE III DUAL USE COMMERCIALIZATION: Military Application: Structurally incorporated lightweight EMP protection for next generation assets that will face hostile electromagnetic environments. Commercial Application: Lightweight protection for avionic and communication systems on board commercial aircraft from personal electronic devices and stray radar energy.
1. Department of Defense Interface Standard, “Electromagnetic Environmental Effects Requirements for Systems,” MIL-STD-464C, Retrieved from http://www.everyspec.com/MIL-STD/MIL-STD+(0300+-+0499)/MIL-STD-464C_28312/, 1 December 2010. 2. Department of Defense Military Standard, “High-Altitude Electromagnetic Pulse (HEMP) Protection for Ground-Based C4I Facilities Performing Critical, Time-Urgent Missions—Vol. 1: Fixed Facilities”, MIL-STD-188-125A, Retrieved from https://assist.daps.dla.mil/docimages/A/0000/0007/1249/000000014229_000000173722_YRNNPHFXPP.PDF?CFID=27815114&CFTOKEN=94644074&jsessionid=5c30fabe1b4251daae1a4d6e554a3e2a7722, 15 February 1994. 3. “Report of the Commission to Assess the Threat to the United States From Electromagnetic Pulse (EMP) Attack—Critical National Infrastructures,” ISBN 978-0-16-080927-9, Retrieved from http://www.empcommission.org/docs/A2473-EMP_Commission-7MB.pdf, April 2008. KEYWORDS: conductive composite, electromagnetic (EM) pulse, EM, electromagnetic interference (EMI) enclosure, EMI, hardening, high-altitude electromagnetic pulse (HEMP), HEMP, radio frequency (RF) shielding, RF AF121-112 TITLE: Near-Surface Residual Stress Measurements for Aerospace Structures TECHNOLOGY AREAS: Air Platform OBJECTIVE: Develop robust methods to measure near-surface residual stresses in complex aerospace structural components. DESCRIPTION: Current design methods for aerospace structure often do not permit the explicit introduction of either bulk or localized residual stresses. Tensile residual stresses are one of many design uncertainties that are accounted for either through the use of conservative material property data or elevated margin of safety requirements. Modern, advanced design methods (e.g., finite-element analysis coupled with fatigue crack initiation and fatigue crack growth analysis) do enable lightweight, sophisticated designs with reduced safety margins; however, the uncertainty associated with residual stress limits the designer’s ability to fully optimize such structures (i.e., residual stress is often treated as a known unknown). Turbine engine and airframe manufacturers are aware of the potentially significant benefits that can arise from explicit accounting for residual stresses (namely, weight reduction in residual stress-free areas and enhanced structural integrity in tensile residual stress areas) and have begun to identify approaches for including residual stress effects in design. To understand and predict the effects of residual stress on fatigue durability or crack initiation (as contrasted with damage tolerance or long crack growth) accurate and reliable residual stress data are required in the near-surface region over depths of roughly 0.025 to 1.25 mm (0.001 to 0.050 inch). Many methods currently exist for the measurement of such near-surface residual stresses; however, few have accuracy and repeatability in the full range of aerospace materials suitable for routine use in correlating fatigue performance with residual stress. Classical x-ray diffraction with layer removal can provide useful data in certain materials and surface conditions, but it also produces noisy and inconsistent data in a number of aerospace metals (e.g., titanium, nickel, and certain aluminum alloys), especially in the presence of large and/or highly textured grains, chemical variations, or multiple phases and precipitates. An advanced method for residual stress measurement is required to further improve reliability of analyses involving residual stress effects. An improved residual stress measurement method would demonstrate the following characteristics: • Applicable to aerospace structural materials (e.g., IN100, Ti-6Al-4V, AA7085, etc.) • Measurement of residual stress to a depth of 1.25 mm (0.050 inch; or greater) • Incremental depth resolution of 0.01 mm (0.0004 inch; or smaller) • Measurement repeatability of ± 10% of the peak stress value (expected to be approximately 10 ksi) over the entire depth range • Measurement accuracy of ± 10% of the peak stress value (demonstrated on a specimen with a well characterized residual stress distribution) over the entire depth range
Military Application: The technology developed will be applicable to the design of more fuel-efficient and durable gas turbine engines and lighter weight unitized airframe structure. Commercial Application: Commercial ships, airliners, and military transports have similar engines and airframes; thus, the technology will be applicable to the design of more fuel-efficient engines and lighter weight airframe structure.
1. D. Ball et al., "Toward Understanding the Impact of Bulk Residual Stress on the Life, Weight and Cost of Primary Aircraft Structure," 2010 Residual Stress Summit, Lake Tahoe CA, Retrieved from: http://sem-proceedings.com/rss4/sem.org-4th-Residual-Stress-Summit-Ball-Toward-Understanding-Impact-Bulk-Residual-Stress-Life-Weight.pdf.
AF121-113 TITLE: Residual Stress Engineering for Aerospace Structural Forgings TECHNOLOGY AREAS: Air Platform OBJECTIVE: Develop modeling and measurement tools to account quantitatively for the effects of bulk residual stresses on machining distortion and the fatigue life of complex aerospace structural components. DESCRIPTION: In an industry-wide effort to reduce weight, increase durability, and reduce cost, traditional built-up aircraft structure and multipiece engine components are being replaced with monolithic structures fabricated from single-piece forgings (e.g., large bulkheads). Such forgings include a wide range of geometries (from bulkheads to integrally bladed rotors) and materials (aluminum, titanium, and nickel alloys). Due to the nature of forging processes (e.g., heat treating, quenching, and cold working), bulk residual stress fields develop during manufacture. These locked-in stress fields subsequently cause problems such as distortion during machining operations and significantly impact fatigue performance. Over the last decade, Integrated Computational Materials Engineering (ICME) methods and tools to quantify bulk residual stress fields have been developed for aerospace forgings. At the same time, new methods have emerged for the measurement of bulk residual stress fields (e.g., the contour method), and these have been successfully applied in aircraft engine and structural forgings. Consequently, for the first time, it is possible to envision a part acquisition/qualification scenario in which the bulk residual stress fields in a finished part are determined by the part vendor and delivered as a data package with each ship set. This residual stress data would become part of the data set used to determine whether or not the part meets the acquiring entity’s (OEM’s) qualification/acceptance criteria and would become part of the quality assurance process for forgings. Developing a quality assurance approach for bulk residual stress management in forgings clearly depends upon the role of the quality data: whether it is used to assure a new production practice (i.e., first-article inspection), assure a group of parts prior to release (i.e., lot-release inspection), or to assure individual part characteristics (i.e., per-article inspection). A quality assurance procedure might use a baseline forging simulation to forecast the nominal residual stress field throughout the component volume. Next, a set of physical measurements would be defined consistent with the role of the quality data (i.e., first-article, lot-release, or per-article inspections). Further forging simulations using known process variabilities would then predict the expected variability of the residual stress field. These forging simulations combined with the residual stress measurement data would become the key elements of a quality assurance program. PHASE I: Demonstrate the proof of concept for a combined bulk residual stress modeling-measurement quality assurance procedure on a first-article forging. With assistance from the TPOC verify relevance and viability of the approach with perspective users. Particular attention should be given in the proposal to the validation protocol of the technique. PHASE II: With assistance from the TPOC fully demonstrate the bulk residual stress quality assurance procedure using production parts. This demonstration should be conducted in a fully relevant production environment. Develop a set of procedures defining a prescribed method of quality management for residual stress in forgings and a strategy for commercialization. PHASE III DUAL USE COMMERCIALIZATION: Military Application: The technology developed will be applicable to the design of more fuel-efficient and durable gas turbine engines and lighter weight unitized airframe structure. Commercial Application: Airliners and military transports have similar engines and airframes; thus, the technology will be applicable to the design of more fuel-efficient engines and lighter weight structure.
1. D. Ball et al., "Toward Understanding the Impact of Bulk Residual Stress on the Life, Weight and Cost of Primary Aircraft Structure," 2010 Residual Stress Summit, Lake Tahoe CA, Retrieved from: http://sem-proceedings.com/rss4/sem.org-4th-Residual-Stress-Summit-Ball-Toward-Understanding-Impact-Bulk-Residual-Stress-Life-Weight.pdf.
AF121-114 TITLE: Lightweight Active Anti-Icing/De-Icing for Remotely Piloted Aircraft (RPA) TECHNOLOGY AREAS: Materials/Processes OBJECTIVE: Develop a lightweight, low power, retrofittable solution for in-flight anti-icing/de-icing for RPAs. DESCRIPTION: RPAs typically do not have any onboard anti-icing/de-icing system(s) to protect critical aircraft surfaces (i.e., wing and tail leading edges, engine inlet lip) from hazardous ice formation. Both commercial and military aircraft typically employ active anti-icing/de-icing systems such as bleed air, bladders, or other pneumatic devices to prevent the ice accumulation. These systems are bulky, heavy, and consume too much power for use on most RPAs. To ensure safety of flight, the FAA regulates flight into known icing conditions and requires an onboard anti-icing/de-icing protection system. FAA requires validation of aircraft anti-icing/de-icing system(s) for Icing Airworthiness Certification. Lacking an in-flight anti-icing/de-icing capability, the operators of unmanned vehicles rely on icing forecasts and onboard icing detectors to avoid icing conditions. Previous research has yielded ice-phobic coatings, which reduced the forces required to remove ice from aircraft wings. However, these have not yet been validated on RPAs during actual flight. Many RPAs employ a laminar airflow design, which requires an ice-free wing in order to meet performance and stability and control requirements. This solicitation requests the design, construction, delivery, and demonstration of a lightweight, low power consumption, reliable, maintainable, retrofittable icing solution for RPA flight critical surfaces (wing, tail, engine inlet lip). An in-flight anti-icing/de-icing capability must be demonstrated for the range of conditions up to 30,000 feet, temperatures from -22 to +32 °F, with a median droplet size of 15 to 40 microns and liquid water content 0.2 to 1.0 gram/cubic meter. This capability can be demonstrated in an icing wind tunnel in accordance with FAA guidance (Ref 1.). PHASE I: Design a laboratory scale concept for in-flight anti-icing/de-icing. Develop a laboratory test plan and conduct screening tests to prove concept for stated droplet size, liquid water content, temperature, altitude, and speed. PHASE II: Construct the active anti-icing/de-icing system developed in Phase I and demonstrate in an icing wind tunnel to stated requirements. During Phase II, identify and partner with an RPA manufacturer to retrofit and demonstrate the technology on a selected airfoil test article. This selected airfoil test article, no larger than 4'' x 6'', will be provided to the contractor by the government. The contractor will need to address cost and schedule for the icing facility testing. The tests performed in the icing facility should replicate conditions for the selected platform and be in accordance with FAA guidance. Systems requiring electrical power consumption should be designed to be compatible with selected RPA platform. PHASE III DUAL USE COMMERCIALIZATION: Military Application: A lightweight, retrofittable icing solution would be of utility to both manned and unmanned small aircraft. Commercial Application: This dual-use technology applies to both military and commercial aircraft concerned with icing. REFERENCES: 1. 14CFR25, Appendix C to Part 25, “Part I—Atmospheric Icing Conditions,” and “Part II—Airframe Ice Accretions for Showing Compliance With Subpart B,” 2010, Retrieved from http://cfr.regstoday.com/14cfr25.aspx 2. FAA Safety Advisory, “Aircraft Icing,” Weather No. 1, 2008, Retrieved from http://www.aopa.org/asf/publications/sa11.pdf 3. Aircraft Icing Handbook, Civil Aviation Authority, Version 1, 2000, Retrieved from http://www.caa.govt.nz/safety_info/GAPs/Aircraft_Icing_Handbook.pdf KEYWORDS: ice accretion, light to moderate icing conditions, remotely piloted aircraft (RPA), RPA, rime ice, runback icing AF121-115 TITLE: Fabrication and Process Optimization of Thick Laminates (= 40 ply) From High- Temperature Polyimide/Carbon Fiber Composites TECHNOLOGY AREAS: Materials/Processes Technology related to this topic is restricted under the International Traffic in Arms Regulation (ITAR) (DFARS 252.204-7009). As such, export-controlled data restrictions apply. Offerors must disclose any proposed use of foreign citizens, including country of origin, type of visa/work permit held, and the Statement of Work (SOW) tasks to be performed. In addition, this acquisition involves technology with military or space application. Therefore, only U.S. contractors registered and certified with the Defense Logistics Services Center (DLSC), Federal Center, Battle Creek MI 49017-3084, (800) 352-3572, are eligible for award. If selected, the firm must submit a copy of an approved DD Form 2345, Militarily Critical Technical Data Agreement. OBJECTIVE: Rapid advancement and optimization of the fabrication and processing of a dual-use polyimide/carbon fiber composite system with high-performance thermal oxidative stability (= 500 to 550°F) for use in the production of turbine engine components. DESCRIPTION: Numerous potential turbine engine applications, such as stators, fan inlet cases, ducts, and external flaps, exist for high-temperature, polyimide matrix, carbon-fiber-reinforced composites on Department of Defense (DoD) weapon systems. For example, large-acreage, lightly-loaded components such as engine ducts are prime candidates for high-temperature polymer matrix composites as they can reduce weight by 40% over their titanium counterparts. For replacement of existing metallic components with polyimide composites, the new materials must be able to operate at elevated temperatures for thousands of hours. High-temperature polymer matrix composites can meet this objective. However, the lack of robust cure cycles and inconsistent process fabrication prevents thick steps, flanges, or pad-ups from being exploited in design concepts. In fact, in some instances, part quality, variability, and defects can result in high production scrap rates well in excess of 20% for large finished parts, which can individually cost as much as $100K. To support emerging production capability of composite engine components and enhance further implementation of high-temperature composites, improvements in manufacturing yield and cost must be demonstrated. The program goal is to improve high-temperature polyimide processing for future realization in manufacturing. The rapid processing advancement is required for a dual-use (military and commercial) polyimide system that can be manufactured for thick sections (= 40 ply) with high-performance thermal oxidative stability (= 500 to 550°F). The polyimide system must demonstrate high-temperature performance and durability equivalent to or better than PMR-15 but without the use of carcinogenic monomers. It is desired that the polyimide/fiber system be somewhat mature (i.e., Technical Readiness Level = 3 to 4) upon award of the contract. It is preferred that the system be dual-use (military and commercial) to reduce cost to the US Air Force. Proposals utilizing optimized cure cycles for current commercially available polyimide systems will be accepted. The cure cycle and bagging scheme of the dual-use material must be developed and successfully demonstrated to routinely result in aerospace quality (< 20% scrap rate, < 2% voids) carbon-fiber reinforced, 40-ply, 12 in. x 12 in., laminated panels (Phase I). Representative duct subelements, such as flanges, access holes, steps, and pad -ups must also be demonstrated (Phase II). The system must retain statistically equivalent mechanical properties at ambient and 550°F for at least 1,000 hours in a dry environment. In Phase I, a preliminary mechanical screening matrix, including uniaxial tension, 4-pt. flexure, compression, and glass transition temperature (Tg) testing according to American Society for Testing and Materials (ASTM) methods of 12-ply laminates, will need to be performed. Mechanical testing at ambient and elevated temperature (550°F) and Tg testing at moisture saturated (wet) conditions is also expected to demonstrate the basic material properties. In Phase II, representative subelements, based on geometries of interest provided by the government, should be manufactured with the improved process. The part quality of the subelements should be evaluated for curved beam strength with the appropriate test specimen geometry and sampling method as defined in ASTM D 6415. Resin kinetic and viscosity material information and real-time data acquired during optimized cure cycle should be provided. In addition, a preliminary assessment of the yield and cost savings should be projected for the improved process. PHASE I: Develop robust manufacturing process concepts via fabrication of a 40-ply, 12 in. x 12 in., polyimide/T650 35 carbon laminate with < 2% voids as evidenced by C-scan and photomicrographs. Perform mechanical screening tests to establish basic properties and demonstrate hot performance (= 500 to 550°F use for 1,000 hours) using ASTM test methods. PHASE II: Advance effort to fabrication of larger parts and representative flange or duct subelements per ASTM D 6415 (same performance metrics as Phase I). Further characterize kinetics, viscosity, mechanical, and physical properties through in-situ monitoring and other analysis methods. Initial analytical material model development (kinetics/viscosity) is desired. Provide a preliminary assessment of the yield and cost savings for the improved process. PHASE III DUAL USE COMMERCIALIZATION: Military Application: Any turbine engine applications utilizing high-temperature, polyimide-matrix, carbon-fiber-reinforced composites on DoD weapon systems. Commercial Application: Cost-competitive, dual-use polyimide/fiber system for commercial aircraft engine components, automobile engine components, etc.
1. Robert A. Gray, “Improved Manufacturing Technologies for Polymer Matrix Composite Engine Components,” Maverick Corporation, AFRL-ML-WP-TR-2007-4028, Accession Nos. DF666665 and ADB326128, Department of the Air Force, April 2007. KEYWORDS: composites, manufacturing, polyimides, processing AF121-120 TITLE: Surface Preparation of Organic Matrix Composites (OMCs) for Structural Adhesive Bonding TECHNOLOGY AREAS: Materials/Processes Technology related to this topic is restricted under the International Traffic in Arms Regulation (ITAR) (DFARS 252.204-7009). As such, export-controlled data restrictions apply. Offerors must disclose any proposed use of foreign citizens, including country of origin, type of visa/work permit held, and the Statement of Work (SOW) tasks to be performed. In addition, this acquisition involves technology with military or space application. Therefore, only U.S. contractors registered and certified with the Defense Logistics Services Center (DLSC), Federal Center, Battle Creek MI 49017-3084, (800) 352-3572, are eligible for award. If selected, the firm must submit a copy of an approved DD Form 2345, Militarily Critical Technical Data Agreement. Directory: osbp -> sbir -> solicitations -> sbir20121 solicitations -> Navy sbir fy09. 1 Proposal submission instructions solicitations -> Army 16. 3 Small Business Innovation Research (sbir) Proposal Submission Instructions solicitations -> Army 14. 1 Small Business Innovation Research (sbir) Proposal Submission Instructions solicitations -> Navy small business innovation research program submitting Proposals on Navy Topics solicitations -> Navy small business innovation research program solicitations -> Armament research, development and engineering center solicitations -> Army 17. 1 Small Business Innovation Research (sbir) Proposal Submission Instructions solicitations -> Navy 11. 3 Small Business Innovation Research (sbir) Proposal Submission Instructions solicitations -> Department of the navy (don) 16. 2 Small Business Innovation Research (sbir) Proposal Submission Instructions introduction sbir20121 -> Air force 12. 1 Small Business Innovation Research (sbir) Proposal Submission Instructions Download 1.28 Mb. Share with your friends:
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