Air force 14. 1 Small Business Innovation Research (sbir) Proposal Submission Instructions
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KEY TECHNOLOGY AREA(S): Air Platforms OBJECTIVE: Develop and demonstrate a smart conceptual design tool to enable improved estimates of performance, including weight and balance associated with early definition of subsystems layout and integration. DESCRIPTION: A conceptual design is characterized as authentic (a closed design) when it includes an authentic development process to define primary parameters that predict the aircraft’s major components and attributes: airframe (could be manufactured), it’s aerodynamic shape (promises flight worthiness), and its propulsion (performance could be achieved). In such a development process, basic engineering and technology assumptions are made. These must be evaluated by further rigorous analysis and/or limited experiments (such as wind tunnel test). This work either verifies the goodness of the design or directs the need for change; specialists are part of the team. When design satisfaction is achieved (low technical risk), a much larger design-manufacturing-administrative team, may then be assigned-and their associated large support funding. This smart conceptual design approach shall differ from the major aircraft industry; it will not likely be able to afford major technical computational and 3-D graphics tools (CATIA, Unigraphics, etc.) and the dedicated individuals who must retain full proficiency in their use. It also differs in that the small team designer(s) must have broader expertise in aeronautical science than one who is tasked only with complex graphical configuration development. The usual fallback from these advanced tools, for the limited aircraft industry, is the drafting board. It provides integration; the external aircraft shape (a three-view drawing) with the internal subsystems (an inboard profile) in one area-on the board. However, this 2-D approach is one to be superseded with a smart semi-automation tool, possibly based on a generic algorithm to cut through the ambiguity of tentative requirements, and to produce a prioritized range of candidate concepts. This part of the tool set is to allow quick development of multiple design choices by a small team or single designer. The primary conceptual design parameters for either the major or limited companies are: the external aerodynamic shape/volume, the speed-altitude region of operation and the associated internal and structural architecture. The architecture of the subsystems is defined by their weight/shape/function/location/servicing and their impact on the flight balance/inertia of the overall design. The minimum, in weight definition terms, are: the structural group, the propulsion group, and the equipment group. These then sum to be the empty weight of the vehicle. The payload, fuel, and crew (if any) are added to this empty weight to define the final Normal Takeoff Gross Weight (NTOGW). The definition (the design) of the internal airframe subsystems includes: installation, strength, thermal environment, power and other special considerations. The engine installation, landing gear arrangement, and external packages also require attention. The weight of the smaller child-components of the subsystems (the sub-subsystems) shall be included in their parent subsystem weight, and their volume allowance must be included in the outer aerodynamic shape’s volume. All of this defining design data will be described in lists, diagrams, tables, and with a 3-D configuration graphic(s). The latter shall show a colored rendering and also translucent views showing the major subsystems. The design is checked for stability, control, and aero-elasticity integrity. It may also be exposed to technology trades, requirement trades, configuration trades, concept comparisons, operational simulations or cost estimation. The conceptual design tool-set must be flexible (probably modular) to handle different classes of aircraft, and nonconventional arrangements. When a final design is obtained (fixed), a larger group of tech specialists takes over for the detail design and manufacturing effort. PHASE I: Demonstrate the feasibility of an overall fixed wing smart conceptual design tool; it may contain new or modified sub-tools. It shall avoid high maintenance (dedicated human proficiency) needs, and its 3-D parametric graphics integration task must bypass the simple 2-D drafting board approach. The development of an intuitive graphics approach is a critical goal of Phase I. PHASE II: This phase will follow through to develop a smooth integration of the sub-tools and associated data bases (e.g., subsystems) that support a fully developed ready to go computerized conceptual design tool. Several classes of military aircraft shall be demonstrated (fighters, transports, bombers). A user instruction manual shall be a part of the final report. Hypersonic, airships and rotary wing design cases are not included in the present scope of the effort. PHASE III DUAL USE APPLICATIONS: Applications include USAF and other service aircraft, and also commercial and general aviation aircraft. The tool shall be able to handle new designs or legacy aircraft modifications. The tool should be modular to enable learning: technology variations, special design cases, updates in procedures. REFERENCES: 1. Nicolai, L.M., and Carichner, G.E., "Fundamentals of Aircraft and Airship Design, Volume I – Aircraft Design," AIAA Educational Series, 2010 Version. 2. Vehicle Sketch Pad, http://openvsp.org/blogs/announcements/2012/12/08. 3. “Weight and Balance Data Reporting Forms for Aircraft, Part I Group Weights Statement," Mil-Std-1374A (Form 380-U-1). 4. Smith, H.K., and Burnham, R., "The Outside has to be Bigger than the Inside," AIAA-80-0726. 5. Waterman D.A., "A Guide to Expert Systems," Addison-Wesley Publishing Company, 1986. KEYWORDS: fixed wing aircraft, conceptual design, aerodynamics, structural airframe, propulsion integration, multidisciplinary design optimization (MDO) AF141-084 TITLE: Radiation Model Development for Combustion Systems KEY TECHNOLOGY AREA(S): Air Platforms OBJECTIVE: Development of physics-based engineering models and corresponding validation procedures as well as associated modules/libraries for radiation heat transfer prediction in combustion systems of relevance to the Air Force (AF). DESCRIPTION: Advanced physics-based modeling and simulation (M&S) tools are playing an increasingly important role in the design of high-performing combustion devices for AF propulsion systems such as rockets, gas turbines, and scramjet combustors. The prediction of engine performance, reliability, and lifetime typically requires the use of coupled multiphysics models. Among relevant physical phenomena, radiation becomes increasingly important at elevated engine working temperature and pressure, especially applicable to large-scale scramjet (e.g., 10X) and high-pressure liquid rockets. Two major impact areas of radiation are 1) combustion process that is crucial to performance and 2) thermal management that is critical to structural design and engine weight [1,2]. Inclusion of radiation heat transfer models that are properly coupled with the turbulence and combustion are critical in engine M&S tools. This topic concerns the development of physics-based engineering radiation models for relevant regimes covering high-pressure (>100 atm) and high-speed (supersonic/hypersonic) turbulent reacting flow conditions. This effort will take advantage of recent, state-of-the-art developments in fundamental radiation algorithmic approaches such as those supported by AFOSR and other agencies. The starting point of the effort requires an evaluation of the existing models, including their fundamental assumptions with respect to AF applications. The effort is interested in innovative, sufficiently accurate, and efficient models and procedures for 1) spatial integration of the radiation transfer equation (RTE), such as other spatial integrations based on spherical harmonics, discrete ordinates, zonal methods, Monte Carlo or similar methods [3]; 2) spectral integration dealing with the complex spectrum of hydrocarbon combustion products relevant and important to AF propulsion systems; and 3) coupling between spatial and spectral integrations. The models are expected to operate within a Reynolds-Averaged Navier-Stokes or a large eddy simulations solution procedure. Related approaches that incorporate the effects of turbulence-radiation interactions as well as radiation-chemistry coupling [4] are also of interest. In addition, radiative properties of soot and the effects of carbon deposits on chamber walls are relevant [5]. An additional area of interest is solution procedures to determine the state of the flow from emission and absorption measurements [6]. In all cases, the models should be capable of operating efficiently in a distributed computing environment. Model predictions should be validated versus relevant data being obtained by related AFOSR and/or AFRL programs. Reacting boundary layers with injected liquid and gas hydrocarbon fuel coolant films are of particular interest. Model development should be carried out in a modular fashion through the specification of standardized application programming interfaces (APIs), which would enable the models to be available as plug-in libraries for computation fluid dynamics (CFD) codes of relevance to the AF and its contractors. PHASE I: Evaluate current radiation modeling assumptions and define validation procedure for their respective use in AF applications including regimes associated with high speed and high pressure. Demonstrate the capability of radiation-turbulence-combustion models for AF relevant propulsion systems. Develop a strategy for well- characterized and generalized interfaces to facilitate module integration. PHASE II: Further develop/enhance the radiation transfer model capability, including spectrally varying radiation properties and high-pressure regimes. Perform detailed validation for test cases of relevance to the AF. Demonstrate the modular approach for the models in candidate CFD code or codes of relevance to the DoD. PHASE III DUAL USE APPLICATIONS: Radiation transfer phenomena are of key relevance to the performance of military propulsion systems, including rockets and gas turbines, and to nonmilitary systems such as space launch systems, aircraft engines, land-based power systems, and automotive engines. REFERENCES: 1. Liu, J., and Brown, M., "Radiative Heating in Hydrocarbon-Fueled Scramjet," AIAA 2012-3775. 2. Crow, A., Boyd, I., and Terrapon, V., "Radiation Modeling of a Hydrogen Fueled Scramjet," Journal of Thermophysics and Heat Transfer, Vol. 27, pp. 11-21. 3. Modest, M., "Radiation Heat Transfer," Third Edition, Academic Press, 2013. 4. Wu, Y., Haworth, D.C., Modest, M.F., and Cuenot, B., "Direct Numerical Simulation of Turbulence-Radiation Interaction in Premixed Combustion Systems," Proceedings of the Combustion Institute, Vol. 30, 639-646, 2005. 5. Charest, M., Groth, C., and Gulder, O., "Effects of Gravity and Pressure on Laminar Coflow Methane-Air Diffusion Flames at pressures from 1 to 60 atmospheres," Combustion and Flame, 158 (2011), pp. 860-875. 6. Daun, K.J., and Howell, J.R., "Inverse design methods for radiative transfer systems," Journal of Quantitative Spectroscopy and Radiative Transfer, Vol. 93, pp. 43-60, 2005. KEYWORDS: radiation transfer, radiation-turbulence interactions, radiation-chemistry interactions AF141-086 TITLE: Lightweight Detachable Roll Control System KEY TECHNOLOGY AREA(S): Nuclear Technology 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 and validate a lightweight, detachable roll control system design that can be integrated with a solid rocket motor and successfully control the motor’s spin rate. DESCRIPTION: The current Minuteman III (MM III) guidance system requires the vehicle to minimize the spin rate by utilizing a roll control system. The roll control system itself is relatively large and heavy, which influences the design of the overall motor. The current roll control system is only on the Second and Third Stage motors due to the motors having single, fixed nozzles. The First Stage motor has four nozzles, and thus has roll control capability and no need for an additional roll control system. However, the future upgrade/replacement motors for all three stages will most likely contain single flex-seal nozzles due to the direction the Solid Rocket Motor (SRM) industry has progressed. The SRM industry has been driven towards this nozzle configuration to minimize cost and maximize performance while still being able to control thrust vectoring. Additionally, the SRM industry has no high demand need to control the motor spin rate on their commercial motors because the current guidance technologies that they use are not affected by spin rate. The MM III guidance system on the other hand is still affected by spin rate until it can be replaced by an updated guidance system. The MM III guidance system will get upgraded in the future, but it is likely not to occur before the current MM III motors are upgraded/replaced. This means that the new SRMs replacing the current MM III motors will need to still have a roll control system to control the motor’s spin rate. Since the current roll control system is quite bulky, it would be advantageous to have a small, lightweight, simple roll control system for SRMs with single, flex-seal nozzles. It will also help if this system was separate from the SRM and easily removable. This way once the MM III guidance system is upgraded, the roll control system can be taken off from the SRMs already fielded. The process of placing the roll control system on the SRMs can also then be skipped on those motors that have yet to be built or placed in the field. This capability will allow for easier life maintenance and a lower life-cycle cost of the SRMs. PHASE I: Create a conceptual design for a detachable, lightweight roll control system. Demonstrate the feasibility of the conceptual design. Identify the required control authority, mechanical interfaces, components, and overall system architecture and demonstrate feasibility through simulation, analysis, or other means. PHASE II: Refine conceptual design completed in Phase I SBIR, include impacts to current weapon system. Validate design and critical technologies by appropriate methods including but not limited to building a sub-scale prototype along with companion simulation and analysis. PHASE III DUAL USE APPLICATIONS: Detachable, lightweight roll control system could be applicable to rocket motor use for those organizations who wish to go with a less expensive guidance system. The detachable, lightweight roll control system could be used as a replacement in the future upgrade/replacement motors for the MM III. REFERENCES: 1. Brown and Hwang; Introduction to Random Signals and Applied Kalman Filtering, 3rd Ed.; Wiley 1997 2. Kenneth R Britting; Inertial Navigation Systems Analysis, Artech House 2010 3. Anton J. Haug; Bayesian Estimation and Tracking: A Practical Guide; Wiley & Sons, Inc. 2012 4. "Hofmann-Wellenhof, Legat, Wieser; Navigation - Principles of Positioning and Guidance, Chapter 8; Springer 2003" 5. Sutton, G. P., Biblarz, O., "Rocket Propulsion Elements," A Wiley-Interscience Publication, John Wiley & Sons, Now York, Eighth Edition, 2010. 6. Pisacane, V. L., "Fundamentals of Space Systems, John Hopkins University Applied Physicas Laboratorey Series in Science and Engineering, 2nd Edition," Oxford University Press, 2005. 7. Knauber, R. N., "Roll Torques Produced by Fixed-Nozzle solid Rocket Motors," Lockheed Martin Vought Systems, Dallas, Texas, Journal of Spacecraft and Rockets, Vol. 33, No. 6, 1996, pp 789-793. KEYWORDS: Lightweight, Detachable, Roll Control, Solid Rocket Motor, Spin Rate, Flex-seal Nozzle AF141-087 TITLE: Additive manufacturing of Liquid Rocket Engine Components KEY TECHNOLOGY AREA(S): Space Platforms 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 and demonstrate additive manufacturing processes for low-rate production of highly complex liquid rocket engine components. DESCRIPTION: Manufacturing process development for rocket applications poses significant technical challenges due to the low production rate, the high complexity of the parts, and the harsh environments in which the parts must operate. The typical production rate for liquid rocket engines (LREs) is typically in the 10’s per year. This means typical assembly line or casting-type processes with dedicated machinery are likely to be cost inefficient. Additionally, these parts are typically very complex with the need for precision shapes, tight tolerances, and high repeatability in order to ensure identical engine-to-engine performance. Currently, designers must make compromises in the design of components such as impellers and inducers that have highly complex internal flow geometries that would be difficult to repeatedly and cost effectively manufacture at low production rates with current processes. These challenges do not lend themselves to traditional manufacturing options. Various components within the engine can be subject to very low temperatures (as low as -253°C if Liquid Hydrogen propellant is being used) or very high combustion temperatures (>2000°C). Pressures can range up to 10,000 psi.
PHASE I: Demonstrate feasibility & benefits of additive manufacturing in terms of cost, producibility, performance, etc. Design & possibly build 1 or 2 LRE components, including coupons for mechanical & flammability properties testing, with typical materials: Cu-alloys, Ti-6Al-4V, Monel. Discuss surface finish attainable & impacts to performance as well as potential non-destructive inspection techniques. PHASE II: Finalize process development for one or more materials. Confirm process by both fabricating test articles for property determination and prototype LRE component of the same size, shape, and complexity as advanced LRE parts. If possible, test built parts at representative LRE conditions. Deliverables: Report of process development and property validation, detailed plan for fabrication process, prototype LRE component, marketing for Phase II Dual Use Applications, test results if applicable. PHASE III DUAL USE APPLICATIONS: This effort supports current and future DoD space launch applications. It will also support commercial and NASA space launch vehicle development. REFERENCES: 1. G.P. Sutton & O. Biblarz, Rocket Propulsion Elements, 7th Ed., John Wiley & Sons, Inc., New York, 2001, ISBN 0-471-32642-9. 2. D.K. Huzel & D.H. Huang, Modern Engineering for Design of Liquid-Propellant Rocket Engines, Vol 147, Progress in Astronautics and Aeronautics, Published by AIAA, Washington DC., 1992, ISBN 1-56347-013-6. 3. IHPRPT Website: http://www.pr.afrl.af.mil/technology/IHPRPT/ihprpt.html. 4. ASM Handbook: Powder Metal Technologies and Applications, ASM International; 2Rev Ed edition, ISBN: 0-871-70387-4. KEYWORDS: Additive Manufacturing, Selective Laser Manufacturing, Electron Beam Melting, Liquid Rocket Engine AF141-088 TITLE: Lowest Lifecycle Cost (LLC) Expendable Launch Vehicles KEY TECHNOLOGY AREA(S): Space Platforms 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: Identify and validate high payoff LLC propulsion technologies that are applicable to space launch and large ballistic missile vehicle applications. DESCRIPTION: Current expendable launch vehicles are designed using an optimum performance methodology. However, there have been numerous studies and technology demonstrations that suggest designing to minimum cost may have life cycle cost advantages over a performance optimized expendable launch vehicle. Performance optimized launch vehicles tend to maximize their payload mass fraction, or conversely, minimize their Gross Lift-Off Weight (GLOW) for a given payload launched into its intended mission orbit. Performance optimization requires high efficiency propulsion and corresponding high propellant mass fractions. These designs can be complex and have high manufacturing costs. On the other hand, LLC strategies weigh the performance advantages of highly optimized systems with simpler designs that have lower manufacturing costs but a larger GLOW. It should be noted that this SBIR topic is interested in proposals which can support existing National Security Space (NSS) payload weights and orbital requirements which can exceed 25,000 lbs to Geosynchronous Orbit (GEO). Directory: wp-content -> uploads -> 2016 2016 -> 2017 afoCo Landmark Scholarship Program 2016 -> Instructions for the Preparation of Camera-Ready Contributions to the Conference Proceedings 2016 -> Step student Scholarships for Year 10-13 2016 -> The united church of canada 2016 -> Idan raichel biography – May 2016 Brief Producer, keyboardist, Lyricist, composer and Performer Idan Raichel 2016 -> An introduction to centre’s interventions expanding access to justice 2016 -> Sanchar Shakti 2016 -> Unit analysis should help here. We want number of bananas 2016 -> The Autobiography of Elder Joseph Bates 2016 -> The Great Heat and the Rhode Island Deep Water Bay Scallop (Argopecten irradians) Fishery 1875-1905 Download 1.72 Mb. Share with your friends:
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