Submission of proposals
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A02-072 TITLE: Light Weight Material for Ballistic Armor TECHNOLOGY AREAS: Air Platform ACQUISITION PROGRAM: PM, Comanche Helicopter OBJECTIVE: The objective of this effort is to identify promising alternate lightweight ballistic armors and designs for ballistic protection used in rotorcraft, fix wing aircraft, light armour vehicles, high value UAVs, launcher, and launcher platforms. It is known that some ballistic materials are efficient is stopping an armor piercing steel projectile when weight is not a factor. However, current State-of-the-Art designs cannot meet performance requirements given weight constraints of these platforms. An advance in amor State-of-the-Art is desperately needed in order to meet both performance and weight requirements of the developing platforms. The desire is to be able to integrate this technology into current and future Army aircrafts, launcher plateforms, and vehicles. DESCRIPTION: The US Army uses conventional ballistic armor packages such as ceramics, composite matrix, fabric, and metals. This SBIR will look at alternate materials and designs that could be used as a lightweight ballistic armor package for protection of crew, equipments, and weapons. The intent is to increase the ballistic limit performance of current armors packages and to increase understanding on what materials best defeat threat projectiles. PHASE I: Phase I effort will consist of identifying candidate materials. Perform research and analysis on how to process and manufacture the raw materials to obtain high hardness, high strength, and high impact energy. Conduct design studies to determine the maximum thickness, weight, and geometry of the armor package. Model the armor using analytical and hydrocode models to optimize ballistic performance. PHASE II: Raw material purchase, metallurgy, and processing. Manufacture armor packages for ballistic testing. Perform design iterations until the performance of the armor package has a ballistic limit are maximized. The test data will be analyzed to determine the effectiveness of the armor package compared to a baseline. Submit 10 lightweight armor packages to the Government for independent reliability testing. PHASE III: Phase III objectives are to qualify the ballistic armor package design, identify the means to integrating armor package into a rotary, fixed wing aircraft, light armour vehicles, and launcher and launcher platforms. DUAL USE APPLICATIONS: The resulting technology will be applicable to military, commercial aircraft, automotive, and law enforcement industries and any other market require lightweight material to resist fragments or projectile impacting at high velocity. OPERATING AND SUPPORT COST (OSCR) REDUCTION: Direct savings would be recognized by reducing the weight of army weapon systems and would enhance the system performance and capability of the systems. REFERENCES: 1) "Lightweight Ballistic Amor for COMANCHE Helicopter", Oak Ridge National Laborarty, Hansen, J. 2) "Impact of the 7.62 mm APM Projectile Against the Edge of A Metallic Target, SRI, Anderson, C., Chocron, S., and Crosch, D. 3) "Computational Analysis of Lightweight Ballistic Armor For Helicopter, U.S. Army Aviation and Missile Command, McDonald, A. 4) "Impact of Metalic Projectiles on Ceramic Targets: Transtion Between Interface Defeat and Penetration", FOA, Weapons and Protection Division", Lundberg, P., Renstrom, R., and Lundberg, B. KEYWORDS: Ballistic Limit, Analytical model, Hydrocode, Ballistic Protection A02-073 TITLE: High Reduction Ratio Drive System for Small Unmanned Aerial Vehicle (UAV) TECHNOLOGY AREAS: Air Platform OBJECTIVE: The objective of this effort is to develop and demonstrate a high reduction ratio drive system for a small UAV. This speed-reducing device must be lightweight, cost efficient, and have a high reliability. The device must also be no larger than 1ft. in diameter, and have an efficiency greater than 98%. DESCRIPTION: The U. S. military services have utilized small low-cost unmanned air vehicles (UAV’s) for reconnaissance and other important missions. An expanded role for UAV aircraft is projected by DOD sponsored studies and planning activities. Many of the current low-speed UAV aircraft are powered by piston engines of 100 HP or less. Several of these engines drive propellers without the need for a gearbox. These engines burn gasoline, which is not the desired fuel type for DOD vehicles. They also have undesirable vibration characteristics and are difficult to start in cold weather operations. Heavy fuel (JP8) gas turbine engines would make ideal propulsion systems for this type of UAV aircraft. For these reasons, the Army is looking to gas turbine engines that will power small UAV’s. Gas turbine engines present many advantages over piston engines. Not only do they use heavy fuel, but they also weigh significantly less than a piston engine. The problem in using gas turbine engines in small UAV’s is the need for a high reduction ratio drive system. There are current gearbox designs for turboprop engines, but these designs are large and weigh almost as much as the engine itself. Therefore, the Army is seeking innovative techniques to solve this problem. A speed-reducing device to provide a high reduction ratio is needed for a 100 HP or less gas turbine engine, with an input speed in the range of about 100,000 RPM to 150,000 RPM. To provide an acceptable propeller output speed, a reduction ratio in the range of 25:1 to 30:1 is necessary. To meet with Army needs, this speed-reducing device must be lightweight, cost efficient, and have a high reliability. The device must be small (no larger than 1ft. in diameter), and also have an efficiency greater than 98%. PHASE I: The objective of Phase I is to conduct small scale evaluations of potential methods to get a reduction ratio of 25:1. These speed-reducing methods must be durable, reliable, cost efficient, lightweight, and be small in size. The results of these evaluations should identify the potential for each method to produce the desired results and allow selection of those that should be pursued in a Phase II effort. PHASE II: The Phase II objectives are to further develop the selected speed-reducing device for a small UAV. Testing of the selected devices shall be conducted in order to establish the ability of the device. PHASE III: The resulting technology will be applicable to both military and commercial UAV’s. A high reduction ratio drive system could be applicable to drive systems/transmissions of other aircraft as well. REFERENCES: 1) Title: Helicopter Drive System R and M Design Guide. AD Number: ADA069835. Corporate Author: UNITED TECHNOLOGIES COP STARFORD CT SIKORSKY AIRCRAFT DIV. Personal Author: Cormier, K. R.. Report Date: April 01,1979. Media: 91 Pages. Distribution Code: 01 – APPROVED FOR PUBLIC RELEASE. Source Code: 323800. From the Collection: TR42. 2) Assessment of the Harmonic Drive as a High Power 80:1 Speed Reduction Gear Box. AD Number: ADB001538. Corporate Author: ROYAL AIRCRAFT ESTABLISHMENT FARNBOROUGH (ENGLAND). Personal Author: Brighton, D. K. Smith, T. R.. Report Date: April 01,1974. Media: 95 Pages. Distribution Code: 16- DOD AND THEIR CONTRACTORS. Source Code: 310450. From the Collection: TR42. 3) Lewenthal, S. H.: Anderson, N.E.; and Nasvytis, A.L.: Performance of a Nasvytis Multiroller Traction Drive. NASA TP-1378, 1978. 4) Hesse, Walter and Mumford, Nicholas. Jet Propulsion for Aerospace Applications, Second Edition. New York: Pitman Publishing Corporation, 1964. 5) Zucrow, M. Aircraft and Missile Propulsion, Volume II. New York: John Wiley & Sons, Inc., 1958. KEYWORDS: Drive system, UAV, heavy fuel engines, gears A02-074 TITLE: Ultra Wideband Network Datalink TECHNOLOGY AREAS: Information Systems ACQUISITION PROGRAM: PEO Aviation, Tactical Unmanned Aerial Vehicle OBJECTIVE: The objective of this SBIR is to apply Ultra Wideband (UWB) radar technology to develop a wireless network data link capability for Army Unmanned Aerial Vehicle (UAV) and Rotary wing systems. DESCRIPTION: The use of UAVs in military operations will continue to be more demanding. Multiple UAVs will be operating in the same theater alongside manned aircraft such as helicopters and airplanes. A major concern is the lack of an adequate wireless data network to allow effective communication and dissemination of data between airborne platforms in close proximity. UWB technology uses accurate timing of short pulses called “Gaussian monocycles” that spread RF energy across an ultra wideband frequency spectrum. This technology implements a “time hopping” methodology that provides numerous advantages over continuous wave technology. Characteristics of the technology include: l ow power spectral densities (in the microwatt range), immunity to interference, excellent multipath immunity, simple electronics, low power consumption, and high bandwidth. Application of UWB technology as a wireless datalink will significantly enhance the warfighter's ability to communicate and disseminate information in an environment with limited available frequency bandwidth. An advantage of this technology is that it is both lightweight and portable making it suitable for application on all Army aviation assets. Medium Access Control (MAC) wireless LAN standards should be developed to work with the UWB band Physical layer. Existing wireless LAN standards may be used, if possible, to reduce risk. PHASE I: Phase I of this SBIR will entail performing a thorough System Engineering Analysis of applying UWB technology to a UAV Wireless Network Data Link. The first step will be to establish detailed requirements for the datalink system. Trade studies shall be performed to assess the feasibility of applying UWB technology in an Army operational environment. Trade studies should focus on defining UWB range limitations, antenna designs and patterns, link budgets, data framing format, maximum data transfer rates, and packaging issues on Army UAV platforms. Datalink interface requirements shall support an open architecture to allow the system to be hosted on several platforms and allow implementation of a wide variety of sensors and payloads. Results of the trade studies will be used to develop an initial architecture design, including any external systems such as antennas and post processing equipment. The initial design shall be documented in a System Requirements Document. A Phase II hardware design and test plan will be generated to include roadmap for detailed hardware and software design, simulation modeling and hardware demonstration. Existing aircraft models and hardware-in-the-loop (HWIL) test beds shall be utilized in Phase II to minimize cost and schedule. The final product for Phase I will be: 1) Trade Study analyses reports 2) Initial architecture design (hw/sw) 3) Phase II Hardware Design and Test Plan
2) Simulation Model 3) Test Reports and Analysis 4) Proof-Of-Concept Model 5) Detailed Hardware/Software ICDs
REFERENCES: 1) “Ultra Wideband Technology Gains A Boost from New Antennas”, Antenna Systems & Technology, Vol. 4, Issue 1, by Hans Gregory Schantz Ph.D., February, 2001. 2) “Multiple Access with Time-Hopping Impulse Modulation”, R. A. Scholtz, Invited Paper, IEEE MILCOM '93, Boston, MA, Oct. 11-14, 1993. KEYWORDS: Network Datalink, Ultra Wideband Radar, Time Hopping, Pulse Position Modulation, Gaussian Monocycle A02-075 TITLE: Non-Contacting Torque Sensor for Helicopter Tail Rotor Drive Systems. TECHNOLOGY AREAS: Air Platform ACQUISITION PROGRAM: PM, Comanche Helicopter (RAH-66) OBJECTIVE: Develop an affordable, flight worthy, non-contacting torque sensor capable of detecting and recording the rapid transient torque loads experienced in drive systems of advanced attack and scout rotorcraft. DESCRIPTION: The RAH-66 Comanche is an advanced scout/attack rotorcraft currently under development by the U.S. Army. This rotorcraft is designed to have very high levels of maneuverability to increase its lethality and survivability. One of the key technologies that provide this high level of maneuverability is the use of a ducted fan in place of a conventional tail rotor. Flight-testing conducted to date has demonstrated that the fantail device has the potential for high transient torque absorption during certain flight regimes and maneuvers. Typical design practice is for all dynamic components to have infinite fatigue life. Designing the fantail drive system for infinite life accounting for the high transient loads results in an unacceptable weight penalty to the aircraft. This situation dictates the need for full time monitoring of torque levels in order to assess the remaining fatigue life of the dynamic components in the drive system. It is envisioned that the desired torque sensor could be located in several locations with different technical challenges. One possibility is to locate the sensor on or around one of the steel gear shafts inside the gearbox. This would continuously expose the sensor to hot oil in the range of 200-300 F. Another possibility is to locate the sensor on or around a section of the tail drive shaft. These shafts are typically thin walled, large diameter organic composite tubes and operate at 6000 RPM. The transient torque events have a maximum duration of 0.08 seconds. Traditional methods, such as strain gages and slip rings, lack adequate long term durability. Desired characteristics of the sensor system are high durability, noncontacting, minimal effect on rotating imbalance of shaft, high accuracy (+ 1% error), insensitivity to operating temperature, and the ability to record torque history and display remaining fatigue life. It is also highly desirable that all electromagnetic signal be locally contained within the device itself. PHASE I: Develop and conduct a feasibility demonstration of the proposed torque sensor technology. The demonstration shall be conducted on a laboratory scale and shall validate to critical technical challenges associated with the proposed technology. PHASE II: The contrator shall further develop the proposed torque sensor by conducting additional bench testing to fully validate the operating characteristics and develop the associated breadboard electronics necessary to record and display the acquired torque history. The Contractor shall conduct design effort in conjunction with the helicopter manufacturer to allow fabrication and testing of the torque monitoring system on the aircraft. The contractor shall install the device onto an aircraft or appropriate aircraft component test rig and conduct dynamic testing to validate the performance of the device in the aircraft operating environment. PHASE III: Focus on the commercialization of the technology through integration into aircraft manufacturer’s design system for use in current and future development programs. Also pursue application of technology to existing military and commercial rotorcraft drive systems and turbo-shaft engines for improved maintenance capability. DUAL USE APPLICATIONS: The resulting technology will reduce the maintenance costs of rotorcraft drive systems by allowing the tail drive system parts to be utilized to their full fatigue life. The torque sensor technology has numerous applications in ground based vehicular applications involving automotive traction control devices, steering feedback sensors and improved engine performance through reduced torsional vibrations. The application to fixed based rotating machinery for diagnostic purposes is also very large. REFERENCES: 1) SAE Paper 2000-01-0085, "Development of a Magnetoelastic Torque Sensor for Formula 1 and CHAMP Car Racing Applications." 2) ABB Company, Torductor-S torque sensor, http://www.abb.com 3) Parkinson, James R., “Evolution and Innovation for Shaft Torque and RPM Measurement for the 1990’s and Beyond” , Presented at the National Technical Specialists Meeting of the American Helicopter society, Scotsdale, Arizona, October 1990. KEYWORDS: Torque Sensor, Rotorcraft, Drive Shafts A02-076 TITLE: A Dynamic Rotorcraft Model for the Study of Advanced Maneuver Concepts TECHNOLOGY AREAS: Air Platform ACQUISITION PROGRAM: PM, Comanche Helicopter (RAH66) OBJECTIVE: Dynamic model wind-tunnel testing (e.g., forced or self-excited model oscillation) has been important for the fixed-wing community, but has not been employed in the rotorcraft community as part of the validation or conceptual modeling process. Our poor ability to predict rotor response to unsteady motion or controls (caused by the unsteady, aerodynamic wake interactions) makes it clear that the lack of such a dynamic test capability is a major gap in the rotorcraft development process. The objective of this solicitation is to develop a dynamic rotorcraft model for the measurement of partially constrained rotorcraft maneuvers - as a necessary step toward the understanding of fully unconstrained (i.e., free-flight) flight behavior. DESCRIPTION: The operation of helicopters at high speeds or at high lift (as in maneuvers) is marked by the occurrence of extremely complex rotor flows. In addition, the response of a rotor to rapid control inputs is governed by unsteady wake behavior, about which little is now known. This is complicated by the fact that a rotor blade is an inherently light and flexible structure that is easily deformed. These loads, deformations and wake behaviors determine the control response and speed/maneuver limits of a helicopter, and it is not yet possible to predict these (and design for these) with confidence. New models of the unsteady rotor wake have been developed for analysis of maneuvering conditions; but data for validation is inevitably very limited when (as is currently the case) full-scale flight testing is the only source of data. As a result, rotorcraft are currently designed on the basis of best past experience; analysis does not identify many major problems, and these can only be fixed as they occur in flight development. The risk, long duration and extreme expense of this process limits the ability to develop concepts that go beyond past experience. Wind tunnel testing can mitigate much of this cost/risk, but this is not possible for all flight modes. The only alternative beyond full-scale development flight or partial scale flying models is dynamic wind tunnel model testing. This type of testing usually involves forced rigid body motion such as a pitch or roll oscillation. In other cases, this type of testing uses the model controls to produce a partially constrained maneuver. A good example of the desired capability would be the pitch and roll response of the rotor to a large change in cyclic pitch control at high speed and/or thrust (near stall) conditions. These types of test cases provide a natural building block approach to the full understanding of maneuver limits and comprehensive code validation. PHASE I: The first objective of Phase I would be the detailed analysis of the scaling issues associated with dynamic testing at model scale. Using the results of the scaling study, the primary task would involve the conceptual development and preliminary design for the test apparatus/wind tunnel model system with the capability to measure various performance, airload, and wake parameters during partially constrained maneuvers and forced oscillations. The model must have a dependable control system (one able to accurately perform and repeat specific maneuvers) and be able to support an instrumentation suite that is appropriate to performance, airload and wake testing. The goal of this development would be to identify the unique opportunities of such a test capability and to realize one or more of these in a working prototype. Possible opportunities can include (but are not limited to), advancing blade behavior, stall-induced control loads, vehicle response and wake behaviors under a range of control inputs. Identify the unique, conceptually risky, or limiting elements of this configuration and perform testing, where necessary, to demonstrate feasibility. PHASE II: A system, incorporating all the essential concepts and required capability elements (established in Phase I), will be designed, built and demonstrated in a mode of operation that verifies its essential capability. This capability can be used to demonstrate some basic load trends, such as the effects of advance ratio and normal acceleration (rotor lift) on pitch link loads. Schemes to measure unsteady wake behavior (or wake-empennage interference) would also be of interest. The system will finally be assessed for a limited range of proposed demonstrated capabilities (possibilities include: (1) the ability to perform repeated realistic maneuvers with repeatable results, or/and (2) a demonstrated ability to emulate specific proposed loading, wake or control behaviors known to occur in full-scale or extreme flight conditions) - and for it's potential to ultimately emulate the full range of these behaviors and capabilities. PHASE III DUAL USE APPLICATIONS: This model will be an important development both for DoD rotorcraft organizations and for the rotorcraft industry as it will reduce a major cost barrier to the development of new rotor concepts. It will also permit the ability to obtain rotor aeromechanics data under maneuvering flight conditions - an effective impossibility with current wind tunnel models. It is for this reason that the extensive full-scale flight testing acquired the UH-60 Airloads Database. A primary first application of this model will be to perform detailed comparisons with some of these UH-60 maneuver loads and other flight tests. The development of this capability would provide an important first step ultimately in the direction of testing free-flight models in unconstrained maneuvers. This system will enable the Army to better understand maneuver capabilities of its rotorcraft and will enable U.S. industry to improve and design more capable rotorcraft. OPERATING AND SUPPORT COST (OSCR) REDUCTION: Military rotorcraft encounter violent maneuvers whose stall-induced loads quickly consume the allowable fatigue life and require replacement of dynamic components. The requested ability to simulate these loads with inexpensive, easy-to-replace models will enable their minimization through an improved understanding of the loading phenomena and the development of advanced rotor designs. The resulting design advances would reduce operation cost via reduced parts replacement and aircraft loss. REFERENCES: 1) "Aerodynamic and Dynamic Rotary Wing Model Testing in Wind Tunnels and Other Facilities," Harris, F., AGARD Lecture Series no. 63, on Helicopter Aerodynamics and Dynamics, 1973. 2) " Flight Testing of the UH-60A Airloads Aircraft," Kufeld, R.M., Balough, D. L., Cross, J. L., Studebaker, K. F., Jennison, C. D., Bousman, W. G., American Helicopter Society 50th Annual Forum Proceedings, Washington D.C., May 1994. 3) "A Qualitative Examination of Dynamic Stall from Flight Test Data," Bousman, W. G., 53rd Annual Forum of the American Helicopter Society, Virginia Beach, Virginia, April 29 - May .1, 1997. KEYWORDS: rotors, dynamic stall, transonic flow, load prediction, advanced rotor development, rotor model development, wind tunnel rotor models, free-flight rotor models, rotor wake testing, model rotorcraft, dynamic stability testing, forced oscillation Directory: osbp -> SBIR -> solicitations solicitations -> Army sbir 09. 1 Proposal submission instructions dod small Business Innovation (sbir) Program solicitations -> Navy sbir fy09. 1 Proposal submission instructions solicitations -> Army 16. 3 Small Business Innovation Research (sbir) Proposal Submission Instructions solicitations -> Air force 12. 1 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 Download 1.78 Mb. 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