AN INVESTIGATION OF A SCRAMJET CAVITY FLAMEHOLDER USING STEREOSCOPIC PARTICLE IMAGE VELOCIMETRY. Justin Kirik University of Virginia

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1 AN INVESTIGATION OF A SCRAMJET CAVITY FLAMEHOLDER USING STEREOSCOPIC PARTICLE IMAGE VELOCIMETRY Justin Kirik University of Virginia A particle image velocimetry (PIV) investigation of a scramjet combustor with a cavity flameholder is described. The first phase of planned research will apply two-dimensional PIV to a measurement plane within the cavity flameholder and parallel to the flowpath central axis. The resulting data are expected to reveal the interaction of cavity flow structures with the shear layer separating the cavity and main-duct flows. Additionally, recirculation regions within the cavity will be quantified. A second phase of the investigation will provide a cross-sectional view of the combustor velocity field by applying three-dimensional PIV to measurement planes perpendicular to the flowpath central axis. These data will be compared with results from different experimental diagnostics developed by partner institutions. Preparations to date have included the design of new hardware to accommodate unique requirements of this research campaign, and testing is expected to begin in the spring of this year. Introduction Over the course of more than a century of development, aviation performance has been defined by few measures more important than speed. Rapid advances were made in the decades after the success of the Wright brothers in achieving practical powered flight: progressively more powerful piston engines propelled aircraft to speeds well over half that of sound. The development of the gas turbine raised the limits even higher, and by the early 1960s specialized aircraft flew in excess of Mach 3. To travel much faster than this would require an even different type of propulsion, and out of this necessity the ramjet was born. This eliminated the turbomachinery of the jet engine, relying instead on the aircraft s speed to do the work of compressing incoming air. Maximum speeds were increased, but a new limit was approached: beyond Mach 5, the temperature of air after being slowed to the subsonic speeds needed by a ramjet was so great that the constituent molecules began to dissociate. This renders the fundamental thermodynamic model of the ramjet useless, and the obvious answer was to keep airflow within the engine supersonic, using a design known as a supersonic combustion ramjet (scramjet). This has proven exceptionally challenging, as combustion is inherently suited to low-speed flows. Thus the challenge was born of maintaining stable combustion in an environment far removed from previous engineering experience 1-3. The fundamental concept of scramjet propulsion is simple: incoming air is compressed and slowed to low supersonic speeds by a series of shock waves, fuel is injected and burned, and the resulting hot gases are expanded through a nozzle to produce thrust. Of these, combustion is the most difficult of the three steps. Stable zones for flameholding are created in low-speed flows through the addition of obstructions which promote turbulence and recirculation. Such features cannot be used in scramjets for two reasons: obstructions at the flow speeds encountered in scramjets would lead to unacceptable performance losses, and the physical obstructions themselves would be unlikely to survive the high temperatures created by compression of hypersonic (Mach numbers in excess of five) air. A variety of Kirik 1

2 designs have been proposed to address the conflicting requirements of stable flameholding and minimal losses in stagnation pressure, both key measures of performance. One receiving significant research attention is the cavity flameholder, a schematic of which is shown in Fig. 1: Fig. 1: Schematic of laboratory model scramjet combustor section showing cavity flameholder. Flow direction is from left to right. As can be seen in Fig. 1, this design incorporates a trapezoidal recess into one wall of the combustor, which serves to provide a region of recirculation outside of the main duct flow. Combustion in this cavity provides a source of heat and radicals to pilot combustion in the main duct, where most heart release occurs. It is important to distinguish the type of recirculation created by a cavity flameholder from that of a device such as a strut designed to create recirculation in the main duct flow. Stagnation pressure losses induced by a cavity are much smaller since the cavity does not impede the core flow, but interacts with it only to the extent of a shear layer which separates the two 4,5. A key feature of most scramjet designs using cavity flameholders is that they are capable of dual-mode operation, signifying the ability to sustain either subsonic or supersonic combustion. At Mach numbers near five, subsonic combustion is actually preferable, unlike the case for higher Mach numbers. While pure ramjets rely on cross-sectional area constrictions to transition flow from supersonic to subsonic and then back again to supersonic speeds, dual-mode scramjets accomplish these same tasks through aerothermodynamic and not physical constraints. To decelerate the incoming flow to subsonic speeds, a pressure rise is induced by combustion which is sufficiently great to create a series of shock waves upstream of the combustor in a section of the engine termed the isolator. After being slowed below Mach 1 by the shock waves, the flow is then driven back to sonic velocity by the heat release of combustion. Expansion in the nozzle then further accelerates the flow to supersonic speeds, allowing the engine to produce thrust 1. When an aircraft powered by a dualmode scramjet accelerates to speeds between approximately Mach 5 and 10, the flow within the engine naturally transitions to supersonic speeds throughout the entire flowpath, and pure scramjet operation is achieved. The significant research attention to dual-mode scramjets has resulted from this ability to operate over a wide variety of Mach numbers. A key NASA application for scramjet technology is a new generation of reusable launch vehicles for lower-cost and responsive space access. By replacing a lower stage of a launch vehicle with a scramjet-powered aircraft, the significant mass penalty of carrying oxidizer in an equivalent rocketpowered stage is eliminated, reducing overall vehicle mass and cost as a result. It is important to note that a conventional gas turbine or rocket engine is needed to accelerate any scramjet-powered vehicle to the speeds necessary for scramjet operation to begin, and a conventional rocket-powered upper stage would be needed to complete orbital insertion for space-access applications. Since it is believed that scramjets could operate over a range of Mach numbers from approximately 4 to greater than 20, the majority of a space launch flight profile Kirik 2

3 (which extends to Mach 25) would fall in the range of applicability to scramjets. Given the degree to which the conditions of scramjet operation are removed from the bulk of combustion research, experimental characterization has been the cornerstone of research in the field. Conventional wind tunnels expose a model of a complete flight vehicle to representative free-stream conditions, and would be ideal for scramjet testing if not for several key challenges. At the high Mach numbers characteristic of hypersonic flight, the required temperatures and mass flow rates of air are much greater than those in more conventional lower-speed testing. To test anything larger than a small-scale model of a hypersonic aircraft would be prohibitive in terms of the resources required to run such a test facility. The solution to this problem comes in the form of direct-connect testing, which eliminates a whole-vehicle model in a freejet environment and instead supplies air only to the engine flowpath itself, significantly reducing the amount of high-temperature air required. Such a configuration presents relatively few disadvantages relative to the freejet environment, since the inlet and nozzle sections which are neglected in direct-connect testing are the most well-understood and thus least in need of experimental investigation 6. Hypersonic ground test facilities are additionally categorized according to test duration and the means of heating test air. While some facilities compress and heat air in real-time to allow for unlimited test durations, power requirements may be significantly reduced by more slowly filling a reservoir with high-pressure, high-temperature air, and then allowing the reservoir to discharge into the test section. The latter type are termed blow-down facilities, and are much more common than their continuous-flow counterparts due to the aforementioned lower power requirements. With respect to heating air, combusting a fuel is often the most costeffective means, but introduces combustion by-products (vitiates) as a result. These vitiates can have varying effects on the resulting performance of the scramjet engine, and to circumvent this a number of facilities have been constructed which do not introduce such impurities. The means of doing so vary, but common options include electrical resistance heaters and heat exchangers coupled with combustion heaters. Experimental Facility The University of Virginia Supersonic Combustion Facility (UVaSCF) is a directconnect continuous-flow unvitiated tunnel. An electrical resistance heater allows stagnation temperatures of up to 1200 K to be achieved. A modular dual-mode scramjet flowpath has been developed which allows for a variety of optical diagnostics to be applied in a number of flowpath configurations. Work to date has concentrated on operation with a ramp fuel injector, although one measurement campaign has investigated a cavity flameholder 7,8. A computer rendering of the combustor section is seen in Fig. 1. A photograph of the combustor in operation is included in Fig. 2: Fig. 2: Cavity flameholder in operation with ethylene fuel. This photograph clearly shows that combustion is anchored in the cavity flameholder and extends well into the main duct. Ethylene is chosen as a fuel for cavity flameholder experiments, as it is representative of the products of a more Kirik 3

4 complex hydrocarbon after being altered by use as an airframe coolant. Such use of fuel as coolant solves two problems: airframe materials must be cooled to survive the high temperatures of hypersonic flight, and large hydrocarbons must be broken down in order to produce molecules which will react sufficiently fast in the scramjet combustor. Complex hydrocarbons are selected as raw fuels due to their relatively simple logistical requirements for storage onboard a flight vehicle. While hydrogen would be more suitable from a purely chemical kinetic point of view, the challenges of carrying such a fuel would outweigh the benefits 2. Pressure data corresponding to test conditions similar to those shown in Fig. 2 are displayed in the flowing plot: Fig. 3: Pressure data in combustor with cavity flameholder. Pressure rise is induced by air injection 7. These data show a pressure rise characteristic of ramjet (subsonic) mode of operation, although it is important to note that in this particular case the pressure rise was induced by additional injection of air into the combustor and not by the heat release of combustion 7. The pressure drop near the downstream limit of data (+10 in.) corresponds to the end of the combustor section, which exhausted to ambient atmospheric conditions. Experimental Technique Particle image velocimetry (PIV) is a minimally-intrusive diagnostic which has been applied to a wide variety of flows. Initially confined to low-speed liquid flows upon its practical introduction in the 1980s, subsequent advances in computer processing technology have expanded its applications to include the high-speed reacting flows of the UVaSCF. PIV relies on the introduction of small seed particles to a flow, which should be selected so as to faithfully follow the flow structures of interest. These particles are illuminated by two successive short-duration laser pulses separated by a known time interval, and photographs are taken at each laser pulse. The result is a pair of images depicting a small displacement of the seed particles, and computer algorithms are then used to determine velocity 9. In the PIV configuration used at the University of Virginia, the laser light is formed into a sheet, producing a plane of velocity data. When a single camera is used, only the velocity components parallel to the measurement plane may be computed, but with a second camera viewing the measurement plane at an angle oblique to that of the first, the out-of-plane component may be measured as well. Such a three-dimensional technique is referred to as stereoscopic particle image velocimetry (SPIV), and was first applied to scramjet flows by researchers at the University of Virginia 10. Experimental Plan: Phase 1 The first phase of my proposed research will apply two-dimensional PIV to a cavity flameholder in a measurement plane perpendicular to the cavity floor. This will yield a cross-sectional view of the cavity velocity field and its interaction with the bounding shear layer. Test conditions will Kirik 4

5 address variations in fuel flow rates as well as transition between ramjet and scramjet mode operation. The resulting data will be useful not only for fundamental physical understanding of cavity flows, but for the validation of numerical models as well, which form a crucial part of the design cycle of scramjetpowered vehicles. This work will be similar to research I participated in at the Air Force Research Laboratory (AFRL) in the summer of , which in turn built on the firstreported PIV investigation of a cavity flameholder 5. The experimental configuration will be similar to that recently used at the UVaSCF to investigate operation with a ramp fuel injector. In this campaign I assisted a colleague in his investigation of the flow of the hydrogen fuel plume and regions downstream 12. The experimental setup is shown below: measurement plane, and were not configured for SPIV operation. The system of lenses was mounted on a traverse to allow for vertical movement of the laser sheet. As can be seen in Fig. 4, the UVaSCF is mounted vertically, a relatively unusual configuration for facilities of its type; however for consistency with prevailing conventions all data and schematics in this paper will be presented in accordance with a horizontal flowpath. Seed particles consisting of 100 nm diameter SiO 2 spheres were introduced to the test section through the fuel plume, allowing regions with fuel or combustion products to be measured. Mean velocity data from one test of this campaign are presented in the following figure: Fig. 5: Mean velocity data representing combusting flow downstream of hydrogen fuel injection in scramjet flowpath. Location of fuel injection is approximately x=-60 mm, y=5 mm. Bulk flow direction is from left to right 12. Fig 4: Experimental configuration of 2D PIV investigation of dual-mode scramjet with ramp fuel injector. Laser light is simulated for purposes of demonstration. The laser beam, which has a circular crosssection after exiting the laser, is directed into the test cell by a mirror, where it is further redirected through four lenses which form the beam into a sheet with an elliptical crosssection. Two cameras were used to expand the The fuel plume has an initial velocity higher than that of the main duct flow, and turbulent mixing gradually eliminates this disparity as fuel convects downstream of the fuel injector. While seeding of the main duct flow was attempted, particle accumulation on test section windows and insufficient signal from particle reflections prevented useful data from being acquired. While the optical configuration for my upcoming research will be nearly identical to that shown in Fig. 4, the seeding configuration will be changed significantly. Seed particles will not be introduced through the fuel stream, but rather through a port in the flowpath wall upstream of the cavity flameholder. By seeding the incoming boundary layer, seed Kirik 5

6 particles will become entrained in the cavity through mixing in the shear layer separating the cavity and main duct flow. This will allow measurement of the entire cavity volume as well as much of the shear layer above it. Measurements of the boundary layer upstream and downstream of the cavity may also be made. It is important to note a distinction in terminology at this point: a boundary layer occurs at the interface of a solid surface and a fluid, while a shear layer separates two fluid streams. A schematic of the experimental configuration is given in Fig. 6: Fig. 6: Schematic of experimental configuration for first phase of research. Laser sheet is directed into the cavity from above, and measurement plane is viewed from the side. As shown in Figs. 4 and 6, the laser sheet enters the combustor through a window in the wall opposite the cavity flameholder, terminating at a right angle to the cavity-side wall. A camera views the illuminated plane orthogonally through a window in the combustor sidewall. The laser sheet may be moved along the tunnel central axis (left-toright in Fig. 6) to allow investigation of the boundary layer upstream and downstream of the cavity. The experimental configuration for this first phase of research will be similar to that used in my work at AFRL, though the test plan will be significantly different. While the AFRL work focused on the effects of flow distortion 11, work at the University of Virginia will concentrate on the effects of varying fueling rates and locations on combustor performance. Key flowfield features expected to be observed in this measurement campaign include recirculation regions occupying most of the cross-sectional plane. These recirculation regions provide much greater residence time for fuel-air mixture than that encountered in the main duct flow, allowing stable combustion to be maintained. Additionally, the shear layer and its possibly unsteady interaction with the cavity flow will be quantified. Experimental Plan: Phase 2 The second phase of research will focus more generally on the performance of the combustor as a whole. SPIV will be applied in measurement planes perpendicular to the tunnel central axis, producing a crosssectional view of the velocity field within the duct. Data will be taken at approximately four measurement planes, which will also be examined with other experimental diagnostics measuring temperature and chemical species concentration by researchers at partner institutions. Such collaborative research is organized under the National Center for Hypersonic Combined Cycle Propulsion (NCHCCP), which includes eight member universities in addition to industry and government partners. I assisted colleagues in two previous SPIV investigations of a scramjet flowpath incorporating a ramp fuel injector, and my research will extend this to a cavity-based combustor. By providing a more global view of combustor operation than the first phase, this research will allow for the Kirik 6

7 determination of the effects of a number of variables on overall performance. In the past year a colleague and I applied SPIV to several planes within a scramjet flowpath configured with a ramp fuel injector. Example data collected at the exit plane of the flowpath are presented in Fig. 7: Fig. 7: SPIV mean velocity data acquired at the exit of a scramjet flowpath. Flow direction is out of page 12. Flow asymmetry is noticeable in these data, which is not uncommon in experimental facilities of this type. This asymmetry has been documented with other experimental techniques 10, and with proper quantification does not detract from its value in validation of numerical models. Data were only acquired in an irregularly-shaped region due to the fact that as in the previously-presented 2D PIV data, only the fuel stream was seeded, thus limiting data collection to regions of combustion byproducts. My planned SPIV work will circumvent this limitation by seeding the entire duct flow through seed injection in the facility plenum. Quantities measured will include root-mean-square (RMS) velocity in addition to mean velocity, allowing quantification of the turbulence level within the scramjet flowpath. The continuousflow capability of the UVaSCF enables the acquisition of large data sets, and it is anticipated that approximately image pairs will be acquired at each test condition to ensure adequate convergence of measurements. While mean velocity can generally be determined with approximately 100 PIV image pairs, RMS velocity requires approximately an order-of-magnitude more to achieve similar convergence. It is important to note that not every PIV image pair produces a velocity vector at a particular location, and it is the number of vectors, not image pairs, that determines sample size. Experimental Plan: Current Progress The first phase of my research is scheduled to begin testing in the late spring of this year and extend into the summer. Preparations have been underway for the past several months, during which time I have also assisted a colleague in the previouslydiscussed PIV experiments. The bulk of preparatory work has involved facility modifications to enable boundary-layer seeding as depicted schematically in Fig. 6. An angled port will be added to an adapter plate between the isolator and combustor sections of the modular scramjet flowpath tested in the UVaSCF. A computer rendering of this part is included in the following figure: Fig. 8: Computer rendering of adapter plate incorporating seed particle injection port. Exit of seed port on tunnel flowpath surface is marked with an arrow. Kirik 7

8 Apart from the addition of a seed injection port, this adapter plate is otherwise identical to a unit currently installed on the UVaSCF. No tunnel dimensions are changed as a result of this modification, thus enabling direct comparison between the ramp and cavity flameholder configurations. In order to deliver seed particles to the boundary layer as described, they will first be levitated using unheated tunnel air in a fluidized bed seeder, depicted schematically in Fig. 9: Summary Preparations are underway to apply particle image velocimetry measurements to the cavity flameholder of a laboratory-model scramjet combustor. Such measurements will provide cross-sectional views of the cavity velocity field, allowing the influence of factors such as varying fueling rates and locations to be determined. This work will be followed by a second phase of research which will take stereoscopic particle image velocimetry measurements across the crosssection of the entire combustor, providing a more global view of combustor performance. By combining these data with measurements from other experimental techniques, a comprehensive database will be built, enhancing both fundamental physical understanding of cavity flows as well as validation of numerical models. Acknowledgement Fig. 9: Schematic of fluidized bed seeder to be used in first phase of measurements. Air is delivered through the porous plug at the bottom of the fluidizing chamber, which then levitates the seed particles and gradually drives them to the shearing nozzle at the top of the chamber. This nozzle reduces particle agglomeration before exiting the seeder. Metal tubing then directs the seed-air mixture to the injection location. A similar seeder will be used for full-duct seeding in the second phase of measurements, although this design does not use a shearing nozzle. This second phase is scheduled to begin in the fall of this year and conclude in the spring of A portion of personnel support was sponsored by the National Center for Hypersonic Combined Cycle Propulsion grant FA The technical monitors on the grant are Chiping Li (AFOSR) and Aaron Auslender and Rick Gaffney (NASA). Additional support was provided by the Department of Defense (DoD) through the National Defense Science & Engineering Graduate Fellowship (NDSEG) Program. The author would like to acknowledge advisor Professor Christopher Goyne of the University of Virginia, as well as Dr. Campbell Carter of the Air Force Research Laboratory for helpful discussions regarding the seeding method for upcoming work. References 1. Heiser, W. H., and Pratt, D. T., Hypersonic Airbreathing Propulsion, Kirik 8

9 AIAA Education Series, AIAA, Washington, DC, Curran, E. T., Scramjet Engines: The First Forty Years, Journal of Propulsion and Power, Vol. 17, No. 6, 2001, pp Billig, F. S., Research on Supersonic Combustion, Journal of Propulsion and Power, Vol. 9, No. 4, 1993, pp Ben-Yakar, A., and Hanson, R. K., Cavity Flame-Holders for Ignition and Flame Stabilization in Scramjets: An Overview, Journal of Propulsion and Power, Vol. 17, No. 4, 2001, pp Tuttle, S. G., Carter, C. D., and Hsu, K. Y., Particle Image Velocimetry in an Isothermal and Exothermic High-Speed Cavity, 50 th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition, 2012, AIAA Tam, C. J., Hsu, K. Y., Hagenmaier, M., and Raffoul, C., Simulations of Inlet Distortion Effects in a Direct-Connect Scramjet Isolator, 47 th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, 2011, AIAA Tatman, B., Experimental Study of Vitiation Effects on Flameholding in a Hydrocarbon Fueled Dual-Mode Scramjet Combustor, master s thesis, University of Virginia, Charlottesville, VA, Tatman, B., Rockwell, R., Goyne, C., McDaniel, J., and Donohue, J., Experimental Study of Vitiation Effects on Flameholding in a Cavity Flameholder, Journal of Propulsion and Power, Vol. 29, No. 2, 2013, pp Raffel, M., Willert, C., Wereley, S., and Kompenhans, J., Particle Image Velocimetry: A Practical Guide, 2nd Ed., Springer-Verlag, Berlin, Smith, C. T., and Goyne, C. P., Application of Stereoscopic Particle Image Velocimetry to a Dual-Mode Scramjet, Journal of Propulsion and Power, Vol. 27, No. 6, 2011, pp Kirik, J., Goyne, C., Peltier, S., Carter, C., and Hagenmaier, M., Velocimetry Measurements of a Scramjet Cavity Flameholder with Inlet Distortion, 49 th AIAA/ASME/SAE/ASEE Joint Propulsion Conference (JPC), 2013 (accepted for publication). 12. Rice, B., Stereoscopic Particle Image Velocimetry Measurements in a Dual- Mode Scramjet Combustor, Virginia Space Grant Consortium Student Research Conference, Virginia Space Grant Consortium, Norfolk, VA, Kirik 9