While not the only application for the desired technologies, the ARSFSS provides an excellent test-bed environment with realistic fuel system conditions that should be targeted in the research and development process. Specifically, nominal conditions for the fuel flow are: 15-170 lb/hr flow rate through 1/4 inch OD and 3/8 inch OD tubing, at fuel temperatures between 250-500 degrees F, and with pressures between 400-500 psia. Note: experiments are initiated with fuel temperatures near ambient with such conditions potentially serving as baseline conditions. Measurements should not change the composition or properties of the main fuel flow. However, diversion of a very small fraction of the flow (0.1 percent of the total) for sampling could be considered. Solutions that rely upon optical access to the flowing fuel are also acceptable provided that the optical interface developed is compatible with the nominal operating conditions of the ARSFSS as noted above. Desired time resolution for the measurements is one fuel composition measurement per minute or faster.
At this stage of research and development, size and mass of the diagnostic system is not a critical parameter; however, future application to actual aircraft systems will have strict size and weight requirements. Technologies that have no chance of miniaturization, while not excluded, will be considered less likely to support long range goals.
Offerors may request to participate in the installation and/or demonstration of prototype hardware in the ARSFSS located at Wright-Patterson AFB. The installation and demonstration of prototype hardware in the ARSFSS will be facilitated by AFRL personnel, and there is no charge to proposers for ARSFSS operations during the Phase II effort.
PHASE I: Phase I is a proof of concept demonstration of advanced diagnostic techniques and analytical instrumentation to measure the composition or specification of thermally-stressed kerosene-based fuels. Demonstration of the concept at elevated temperatures and pressures that are relevant to next generation air platforms is expected (e.g., fuel flows at 250 to 500 degrees F and 400 to 500 psia).
PHASE II: Phase II is a further development with a prototype demonstration of the diagnostic under realistic ARSFSS or high-performance military aircraft fuel system conditions. Deliverables include the instrumentation developed, schematics, operating manuals and instructions, and report documentation for the overall research effort. Demonstration of the prototype’s capabilities in other areas of application that suggest successful application within an aircraft fuel system would be considered.
PHASE III DUAL USE APPLICATIONS: Phase III is a full-scale prototype demonstration of the advanced diagnostic techniques and analytical instrumentation. Ground demonstration of the full scale prototype in a Government/industry facility that is used to research and develop aircraft subsystem components is expected.
1. Robert W. Morris, Jr., et.al., "Evaluation of the Impact of Kerojet Aquarious Water Scavenger Additive on the Thermal Stability of Jet A Fuels," Defense Technical Information Center, AFRL-RQ-WP-TR-2015-0014.
2. Nicholas J. Kuprowicz, et.al., "Use of Measured Species Class Concentrations with Chemical Kinetic Modeling for the Prediction of Autoxidation and Deposition of Jet Fuels," Energy & Fuels, Vol. 21, pp. 530-544 (2007).
3. Tim Edwards and Steven Zabarnick, "Supercritical Fuel Deposition Mechanisms," Ind. Eng. Chem. Res., Vol. 32(12), pp. 3117-3122 (1993).
KEYWORDS: fuels, thermal management, diagnostics, online, composition
TITLE: Durable Pre-cooling Heat Exchangers for High Mach Flight
TECHNOLOGY AREA(S): Air Platform
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. Gail Nyikon, firstname.lastname@example.org.
OBJECTIVE: Enable turbomachinery operation in excess of Mach 4 by pre-cooling the incoming air. Utilize modern materials, manufacturing, and design processes to design a durable pre-cooling heat exchanger.
DESCRIPTION: Throughout the past 50 years, there have been multiple efforts to develop a single air-breathing engine that is capable of thrust from takeoff up to Mach 4 speeds and faster. A major challenge in developing this capability is the hot air that is ingested at high Mach speeds. There have been attempts dating back to the 1960s to solve this problem by using a pre-cooler heat exchanger in front of the air compressor face to reduce the temperature of the incoming air. It is expected that high Mach flight will be growing in importance for the Air Force to execute its five core missions and that precooled propulsion could be an enabler for new platform capabilities. The objective of this topic is to mature the technology for a lightweight and compact pre-cooler heat exchanger for high Mach propulsion that uses turbomachinery.
No pre-cooling heat exchanger has ever been flown. The biggest difficulty is getting the heat exchanger light weight and compact enough to be practical for flight. In many industries, modern manufacturing has allowed for lighter weight components and unique geometries to be built. This topic will leverage modern manufacturing techniques (e.g., additive manufacturing, friction stir welding, C&C milling, etc.) to develop a pre-cooler heat exchanger that is practical to be used in a propulsion system on a high Mach flight system. At the end of the Phase II, it is expected to fabricate a scaled prototype of the heat exchanger and conduct initial evaluation testing. Throughout this topic, it is important to address thermal integration for the necessary systems involving the heat exchanger.
Important attributes of pre-cooler heat exchangers (of roughly equal importance) that need to be addressed include durability, affordability, ability to integrate with propulsion and flight systems, scalability, manufacturability, impact on ground operations, material and manufacturing maturity, amount of pressure drop across the heat exchanger, and maintainability. The pre-cooler should be able to cool incoming freestream air to about 500 degrees F or cooler for flight conditions at altitudes above 55,000 feet. It is also expected the heat exchanger to be developed will have a specific power of at least 15kW/lbm.
Both Phase I and Phase II will consist of an appropriate level of design and systems engineering efforts to understand what it will take to fully develop the proposed solution. These efforts should address all issues but focus on the demonstrations that will be performed in Phase II. Modeling of the heat exchanger’s performance and its integration is needed throughout both phases to understand its potential. Recommend developing one or more reference vehicle platform designs for one or more Air Force core missions to show how the heat exchanger could enable that capability.
A letter of endorsement from a Versatile Affordable Advanced Turbine Engines (VAATE) participant is highly encouraged.
Commercialization of the pre-cooler heat exchanger involves integration of the pre-cooler into high-speed propulsion systems for DoD and/or commercial needs such as point to point cargo and access to space. Commercialization of the heat exchanger can also be used for propulsion thermal management and terrestrial applications.
Remote access to the DoD Supercomputing Resource Center (DSRC) to cleared personnel will be made available if needed.
PHASE I: Conduct initial design of the pre-cooler heat exchanger with an emphasis on its integration and manufacturing. Based on higher level platform requirements, derive requirements for the heat exchanger components that have early verification and validation. Develop plans for the Phase II fabrication and testing.
PHASE II: Fabricate a scaled prototype of the heat exchanger utilizing the proposed manufacturing approach. Conduct testing in a relevant laboratory environment. Develop and validate performance and lifting models based on the testing. Utilize this information to increase the understanding of how the heat exchanger integrates into a platform or platforms.
PHASE III DUAL USE APPLICATIONS: Phase III will focus on maturing the heat exchanger and beginning to integrate it into a full propulsion system and a vehicle platform. Additional Phase III activities can consist of applying the heat exchanger and its manufacturing to other defense and commercial domains.
1. Murthy, S.N.B., "Developments In High-Speed Vehicle Propulsion Systems," AIAA, 1996.
2. Balepin, V., et.al., "Combined propulsion for SSTO rocket - From conceptual study to demonstrator of deep cooled turbojet," AIAA 96-4497.
3. Taguchi, H., et.al., "Performance Evaluation of Hypersonic Pre-Cooled Turbojet Engine," 20th AIAA International Space Planes and Hypersonic Systems and Technologies Conference, AIAA 2015-3593.
4. Ahren, J.E., "Thermal management of air-breathing propulsion systems," 30th Aerospace Sciences Meeting and Exhibit, AIAA 92-0514, 1992.
5. Murray, J.J., et.al., "An Experimental Precooler for Airbreathing Rocket Engines," 48th International Astronautical Federation Congress Melbourne, Australia 1998, IAF-98-S.5.02.
KEYWORDS: high mach, precooling, precooler, heat exchanger, thermal, materials, manufacturing, design, propulsion, hypersonic
TITLE: Automated Synthesis of Propulsion-Power-Thermal Architectures
TECHNOLOGY AREA(S): Air Platform
OBJECTIVE: Establish a methodology to synthesize innovative architectures to explore and evaluate the full design space of next-generation aerospace propulsion systems. This methodology should highly integrate propulsion, power, and thermal components.
DESCRIPTION: Next-generation military aerospace systems are difficult to design, with requirements (e.g., faster speeds, longer ranges, cost effectiveness) often driving solutions that include compromises for propulsion and vehicle systems. Propulsion architectures have historically been dominated by the traditional assumptions (e.g., turbine engine Brayton cycle). Breaking this trend and developing alternative configurations of propulsion, power, and thermal components may be essential to solve challenges and meet future flight needs. It is hypothesized that alternative configurations of propulsion, power, and thermal components can be formulated; constituting an unconventional architecture that may help solve these challenges (a simple example is inter-turbine burning). The goal of this topic is to develop a software tool to conceptually optimize a propulsion architecture minimizing user assumptions.
Other engineering disciplines, such as the electronic industry, have shown the ability to aggressively seek innovative ideas and concepts through improved architecture development as shown in Refs. 4 and 5. This topic looks to leverage modern computing resources, optimization algorithms, and modeling techniques to synthesize propulsion, power, and thermal architectures, analogous to the electronic industry. At the end of Phase II, a design and analysis software tool for engineers is expected.
Inputs to the synthesis method must include user requirements, design objectives, and constraints. These inputs include the context of the vehicle to be designed. The method should be able to sort through many different combinations of architectures and find which best meet the defined problem. The sorting process should be sufficiently broad to include architecture combinations that are infeasible. It is desired that the method developed under this topic be used in conceptual design studies for future platforms for defense and commercial applications.
A main goal of these methods is to be able to search through architectures that may not be intuitive to designers. This will include trading off different architecture components (e.g., compressors, heat exchangers, etc.) and the way they are laid out. Allowing the user flexibility to define component parameters (e.g. compressor maps) is important to maximize the flexibility. Integration effects with the vehicle shape and other internal systems need to be considered. These can be based on simplified relationships but should allow the user to input additional data and relationships.
Important attributes (of roughly equal importance) of the synthesis method includes the ability to conduct automated trade studies, flexibility with defining a particular problem, portability between operating systems, ability to add in user defined models and data, ease of the graphical user interface, ability to be wrapped in modeling/design/analysis frameworks (e.g., Modelcenter, AML, etc.), and ease of interface to visualization methods.
There are many different propulsion, power, and thermal modeling tools available that could be leveraged for this topic. Some example tools are the Numerical Propulsion System Simulation (NPSS) tool and NASA’s Generalized Fluid System Simulation Program (GFSSP). These current tools require a user to define the propulsion/power/thermal architecture.
The methods developed in this topic are to be put together into a design and analysis software tool. This software should be able to run on a desktop high performance workstation and it is desired to be compatible across platforms (e.g., Windows, Linux). The software should analyze at least 10 different propulsion architectures per minute with an objective of 100 per minute (assuming 2 processors). The software needs to be designed to allow for a user with a background in aerospace propulsion to easily learn to use.
Commercialization opportunities of this topic include using the tool for aerospace design and technology development programs. Other opportunities include commercial sales of the software tool to various aerospace companies, universities, and government agency.
PHASE I: Phase I will survey user needs to develop a set of requirements for the synthesis methods. Using that information, an initial prototype of the synthesis method will be developed followed by initial testing for early requirements validation and verification. To get to an initial prototype, it is expected that some trades studies will be conducted to ensure the best approach is being used.
PHASE II: The objective of Phase II is to have a completed tool based on the developed synthesis methods. This will include using lessons learned from Phase I, further requirements refinement, and testing of the method. The contractor should work with users to develop test cases for various types of military missions.
PHASE III DUAL USE APPLICATIONS: Phase III can consist of improvement to the synthesis methods and use of the methods in Air Force/DoD technology and concept development programs. Phase III can also consist of commercial sales of the tool developed to use in commercial aerospace propulsion development.
1. Evans, A.L., et al., "Numerical Propulsion System Simulation's National Cycle Program," AIAA Paper 98-3113 (1998)
2. Ashleman, Russell H., Jr., Lavelle, Thomas, and Parsons, Frank, "The National Cycle Program: A Flexible System Modeling Architecture," AIAA Paper 98-3114 (1998).
3. Mogavero, A. and Brown, R., "An improved engine analysis and optimization tool for hypersonic combined cycle engines," AIAA 2015-3681, 20th AIAA International Space Planes and Hypersonic Systems and Technologies Conference.
4. Von Moll, Alex and Behbahani, Alireza R., "Comparison of Communication Architectures and Network Topologies for Distributed Propulsion Controls," Presented at the 59th International Instrumentation Symposium, Cleveland OH, ISA 2013, ADA586909 on www.dtic.mil.
5. Chepko, A.B., et al., "A Modeling Framework for Applying Discrete Optimization to System Architecture Selection and Application to In-Situ Resource Utilization," AIAA 2008-6058. 12th AIAA/ISSMO Multidisciplinary Analysis and Optimization Conference 10-12 September 2008, Victoria, British Columbia, Canada.
KEYWORDS: architectures, propulsion, power, thermal, optimization, synthesis
TECHNOLOGY AREA(S): Air Platform
OBJECTIVE: The objective of this research is to accommodate the variability and uncertainty of jet fuel properties during the design of next-generation thermal management systems for tactical aircraft.
DESCRIPTION: Propulsion-integrated thermal management solutions have emerged as a primary driver within the aerospace community. The thermal management needs for next-generation air platforms are generally considered to be an order of magnitude higher than today's advanced systems, and the manner in which fuel is used as a heat sink within air platforms is expected to fundamentally change as the community moves forward.
The underlying reason for this foundational shift reflects a combination of factors that are simultaneously converging: i) at least a 10X increase in the amount of thermal energy content to actively manage; ii) thermal management system approaches that are heavily-dependent on fuels and also require fuel temperatures in excess of 300 degrees F; iii) Air Force utilization of commercial-grade fuels which have a broader specification window than military-grade fuels such as JP-8; and iv) legacy design practices that assume perfect fuels and neglect real-world fuel-to-fuel variability in fuel properties and system performance. This topic addresses factors (iii) and (iv) where current state-of-the-art (SOTA) modeling and simulation tools supporting system-level modeling and performance prediction, often utilized in performance phase-space trade studies, do not currently account for the real world variability of the jet fuel properties. To extend beyond the SOTA, this topic is interested in modeling and simulation tools that are adaptable to the system performance studies, and provide for variability in important fuel properties and performance behaviors. The fuel-to-fuel variation in constant pressure specific heat (Cp) is particularly troublesome from a thermal management perspective, as the measured variability across a range of Jet A fuel samples has been shown to exceed 15 percent. This level of variability directly translates to a 15 percent uncertainty in thermal energy transfers and/or component temperatures in fuel thermal management systems, and at the present time, the community does not account for this variability in the design process. Typical Cp values for a Jet A fuel are 2.0 kJ/kgK at 20 degrees C to 2.65 kJ/kgK at 180 degrees C (Aviation Fuel Properties, Coordinating Research Council, Inc., 1988).
The objective of this research is to accommodate the variability and uncertainty of jet fuel properties during the design of next-generation thermal management systems for tactical aircraft. The development and validation of modeling and simulation capabilities that incorporate key fuel parameters is an integral part of the research effort. Demonstration of the modeling and simulation capabilities to perform design space exploration of conceptual thermal management systems is required. Evaluation of both the static and dynamic responses of highly-integrated thermal management systems to fuel property variation is also required. The establishment of critical design parameters and guidelines for adoption within the community, and in particular the Government/industry partnership of Versatile Affordable Advanced Turbine Engines (VAATE), is strongly desired.
To successfully perform the work described in this topic area, offerors may request to participate in the installation and/or demonstration of prototype hardware in the Advanced Reduced Scale Fuel Systems Simulator (ARSFSS) located at Wright-Patterson Air Force Base, OH. The installation and demonstration of prototype hardware in the ARSFSS will be facilitated by Air Force Research Laboratory (AFRL) personnel. There is no charge for ARSFSS operations during the Phase II effort.
PHASE I: Phase I is a proof-of-concept demonstration of a modeling and simulation approach. The ability to vary any combination of fuel properties (such as constant pressure specific heat) across a range of values within, and exceeding, specification limits is required within a representative thermodynamic model of a tactical aircraft thermal management system.
PHASE II: Phase II is a validation of the modeling and simulation approach. A configuration description and experimental data for the ARSFSS will be provided during Phase II. It is expected that the modeling and simulation approach will be configured to simulate ARSFSS hardware and components, and a comparison of simulated and measured changes in fuel system performance attributable to fuel properties will be performed.
PHASE III DUAL USE APPLICATIONS: Phase III is a validation of the modeling and simulation approach at full scale conditions for both steady state and transient/dynamic modes of operation. The utilization of government/industry facility data that is used to research and develop aircraft subsystem components is expected.
1. Robert W. Morris, Jr., et al., "Evaluation of the Impact of Kerojet Aquarious Water Scavenger Additive on the Thermal Stability of Jet A Fuels," Defense Technical Information Center, AFRL-RQ-WP-TR-2015-0014.
2. Mark Bodie, et al., "Robust Optimization of an Aircraft Power Thermal Management System," 8th Annual International Energy Conversion Engineering Conference, AIAA 2010-7086.
KEYWORDS: thermal, uncertainty, probabilistic, design, fuels, high-temperature
TITLE: Fast Valve for Starting Hypersonic Wind Tunnels
TECHNOLOGY AREA(S): Air Platform
OBJECTIVE: Develop a full-bore valve for air that is capable of opening quickly, sealing well against pressure, and with long fatigue life.
DESCRIPTION: The state of the boundary layer and where and when it transitions, can play a critical role in the survivability and controllability of hypersonic precision strike vehicles. Understanding the physics leading to transition will enable better prediction of boundary layer transition for strike vehicles. This will enhance their survivability and may also reduce the amount of thermal protection needed, thereby increasing the payload mass.
Flight tests are often prohibitively expensive, so the majority of hypersonic boundary-layer transition research is conducted in ground test facilities. To best simulate hypersonic atmospheric flight and study boundary-layer transition in an affordable manner, uniform, hypersonic flow must be established quickly. One increasingly popular design for cost-effective, high-speed, super- and hypersonic wind tunnels is a Ludwieg tube. In this type of facility, a long tube is filled with pressurized gas. This gas is typically isolated from downstream, low-pressure components by means of a burst diaphragm(s) or a valve. To initiate the high-speed flow, the diaphragm(s) is ruptured, or the valve is opened. This allows the high-pressure gas to expand through the nozzle and accelerate to the desired speed. After the gas pressure in the low-pressure end of the tunnel rises sufficiently, the high-speed flow stops and only subsonic flow continues.