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Non-premixed Conditions in the Flameholding Recirculation Region behind a Step in Supersonic Flow


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NON-PREMIXED CONDITIONS IN TH E FLAMEHOLDING RECIRCULATION REGION BEHIND A STEP IN SUPERSONIC FLOW By AMIT THAKUR A DISSERTATION PRESENTED TO THE GRADUATE SCHOOL OF THE UNIVERSITY OF FLOR IDA IN PARTIAL FULFILLMENT OF THE REQUIREMENTS FOR THE DEGREE OF DOCTOR OF PHILOSOPHY UNIVERSITY OF FLORIDA 2006

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Copyright 2006 by AMIT THAKUR

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iii ACKNOWLEDGMENTS I am grateful to my advisor, Dr. Corin Se gal, for guiding me through the research work and for driving me to finish the Ph.D. program in a relatively short time. Working at the combustion laboratory helped me learn about scramjet engines, and provided valuable practical experience ranging fr om setting up, performing and analyzing scientific experiments to writing technical papers and deliveri ng presentations at international conferences. Beyond academics, interaction with fellow lab mates having diverse backgrounds from differe nt countries made for an enjoyable and educational experience in itself. Finally, I thank my mo ther and father for being typical Indian parents, encouraging their child to pursue higher education. This work was performed with support from NASA grant NCC3-994, with Claudia Meyer as the Program Manager.

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iv TABLE OF CONTENTS page ACKNOWLEDGMENTS.................................................................................................iii LIST OF TABLES.............................................................................................................vi LIST OF FIGURES..........................................................................................................vii ABSTRACT.......................................................................................................................ix CHAPTER 1. INTRODUCTION........................................................................................................1 Hypersonic Air-Breathing Vehicle...............................................................................1 Air-breathing Engine Classification.............................................................................2 Turbojet Engine.....................................................................................................2 Ramjet Engine.......................................................................................................3 Scramjet Engine.....................................................................................................4 Flameholding in Supersonic Flow................................................................................5 Cavity Flameholder...............................................................................................8 Strut Flameholder................................................................................................11 Free Shear Layer Fluid Dynamics..............................................................................13 Flame Stability Limits................................................................................................16 Flame Stability in High-Speed Subsonic Flow...................................................16 Non-Premixed Flame Stability in Supersonic Flow............................................19 Species Distribution in NonPremixed Flameholding Region in Supersonic Flow...23 Experiments.........................................................................................................23 Computations.......................................................................................................27 Mass Sampling for Species Concentration Measurement in Supersonic Flow..........27 Optical Diagnostics for Species Concentr ation Measurement in Supersonic Flow...29 Planar Laser-Induced Fluorescence (PLIF).........................................................29 Raman Scattering.................................................................................................33 Scope of Study............................................................................................................34 2. EXPERIMENTAL SETUP........................................................................................39 Supersonic Wind Tunnel............................................................................................39 Mass Sampling and Analysis......................................................................................40 Hardware.............................................................................................................40

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v Data Processing...................................................................................................43 Optical Diagnostics.....................................................................................................44 Hardware.............................................................................................................44 Image Acquisition and Processing......................................................................46 3. RESULTS: NON-REACTING FLOW......................................................................57 Mass Spectrometry (MS)............................................................................................ 57 Base Injection: Helium........................................................................................57 Base Injection: Argon..........................................................................................59 Upstream Injection: Helium................................................................................61 Upstream Injection: Argon..................................................................................62 Planar Laser Induced Fluorescence (PLIF)................................................................. 63 Base Injection: Helium........................................................................................63 Base Injection: Argon..........................................................................................64 Upstream Injection..............................................................................................65 Comparison between MS and PLIF data....................................................................65 4. RESULTS: REACTING FLOW................................................................................78 Base Injection: Hydrogen...........................................................................................78 Upstream Injection: Hydrogen...................................................................................80 5. CONCLUSIONS........................................................................................................88 APPENDIX A. LABVIEW PROGRAM FOR MA SS SAMPLING SEQUENCING........................92 B. MATLAB PROGRAM FOR PLIF IMAGE PROCESSING.....................................99 LIST OF REFERENCES.................................................................................................109 BIOGRAPHICAL SKETCH...........................................................................................115

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vi LIST OF TABLES Table page 2-1. Mass spectrometer calibrati on factors for various gases..........................................50 3-1. Base fuel injection: global and local He..................................................................67 3-2. Base fuel injection: global and local Ar..................................................................67 3-3. Upstream fuel inj ection: global and local He..........................................................67 3-4. Upstream fuel inj ection: global and local Ar..........................................................67 4-1. Base fuel injection: global and local H2.................................................................82

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vii LIST OF FIGURES Figure page 1-1. Comparison between (a) Rocket-powered vehicle [Source: www.isro.org] and (b) Scramjet-powered hypersonic vehicle [Source: www.nasa.gov].......................36 1-2. Schematic diagram of (a) turbojet engine (b) ramjet engine (c) scramjet engine....37 1-3. Various flameholder geometries for supers onic flow: (a) rearward step (b) cavity (c) strut. ................................................................................................................38 1-4. Schematic diagram of supersonic flow over a rearward step...................................38 2-1. Description of the test section show ing fuel injection and mass sampling locations: (a) image and (b) schematic diagram.......................................................51 2-2. Mass sampling from the recirculation region behind the step for analysis by the mass spectrometer: (a) schematic diagram (b) image..............................................52 2-3. MS measurements of species mole frac tion distribution in the recirculation region as the solenoid valves switch se quentially from one sampling port to another under steady experimental cond itions (a) non-reacting flow with helium as the simulant fuel (b) reacting flow with hydrogen combustion...........................53 2-4. Schematic diagram of planar laser -induced fluorescence (PLIF) setup................... 54 2-5. Sample PLIF image for (a) background (b) laser sheet...........................................55 2-6. Temporal variation of laser sheet profile at y /H = 1.1.............................................56 3-1. Wall pressure distribution for non-reacting flow. P0air = 4.8 atm, Mair = 1.6. The axial origin is placed at the step...............................................................................68 3-2. Base fuel injection: MS measurement of helium mole fraction distribution in the recirculation region for (a) wall sampling (b) inflow sampling. P0air = 4.8 atm, Mair = 1.6. ................................................................................................................69 3-3. Base fuel injection: MS measurement of argon mole fraction distribution in the recirculation region for (a) wall sampling (b) inflow sampling. P0air = 4.8 atm, Mair = 1.6. ................................................................................................................70

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viii 3-4. Upstream fuel injection: MS measurem ent of helium mole fraction distribution in the recirculation region for (a) wall sampling (b) inflow sampling. P0air = 4.8 atm, Mair = 1.6..........................................................................................................71 3-5. Upstream fuel injection: MS measuremen t of argon mole fraction distribution in the recirculation region for (a) wall sampling (b) inflow sampling. P0air = 4.8 atm, Mair = 1.6..........................................................................................................72 3-6. PLIF measurement for base inj ection of helium (a) image (b) XHe distribution (%). P0He = 5.4 atm, P0air = 4.8 atm, Mair = 1.6.........................................................73 3-7. PLIF measurement for base inj ection of helium (a) image (b) XHe distribution (%). P0He = 12.0 atm, P0air = 4.8 atm, Mair = 1.6.......................................................74 3-8. PLIF measurement for base in jection of argon (a) image (b) XAr distribution (%). P0Ar = 5.4 atm, P0air = 4.8 atm, Mair = 1.6.................................................................75 3-9. PLIF measurement for base in jection of argon (a) image (b) XAr distribution (%). P0Ar = 12.0 atm, P0air = 4.8 atm, Mair = 1.6...............................................................76 3-10. Comparison between MS and PLIF data for base fuel injection of (a) helium (b) argon. ................................................................................................................77 4-1. Wall pressure distribution for hydrogen combustion tests. P0air = 4.5 atm, Mair = 1.6. ................................................................................................................83 4-2. Base fuel injection: Wall sampling resu lts for (a) hydrogen equivalence ratio distribution in the recirc ulation region (b) combusti on species mole fraction distribution. P0H2 = 4.5 atm, P0air = 4.5 atm, Mair = 1.6............................................84 4-3. Base fuel injection: Wall sampling resu lts for (a) hydrogen equivalence ratio distribution in the recirc ulation region (b) combusti on species mole fraction distribution. P0H2 = 8.2 atm, P0air = 4.5 atm, Mair = 1.6............................................85 4-4. Base fuel injection: Inflow sampling results for (a) hydrogen equivalence ratio distribution in the recirc ulation region (b) combusti on species mole fraction distribution. P0H2 = 4.5 atm, P0air = 4.5 atm, Mair = 1.6............................................86 4-5. Base fuel injection: Inflow sampling results for (a) hydrogen equivalence ratio distribution in the recirc ulation region (b) combusti on species mole fraction distribution. P0H2 = 8.2 atm, P0air = 4.5 atm, Mair = 1.6............................................87

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ix Abstract of Dissertation Pres ented to the Graduate School of the University of Florida in Partial Fulfillment of the Requirements for the Degree of Doctor of Philosophy NON-PREMIXED CONDITIONS IN TH E FLAMEHOLDING RECIRCULATION REGION BEHIND A STEP IN SUPERSONIC FLOW By Amit Thakur May 2006 Chair: Corin Segal Major Department: Mechanical and Aerospace Engineering Flameholding in supersonic flow depends on local conditions in the recirculation region, and on mass transfer into and out of this region. La rge gradients in local gas composition and temperature exist in the recirculation region. Hence, stability parameter correlations developed for premixed flames ca nnot be used to determ ine blowout stability limits for non-premixed flames encountered in practical devices. In the present study, mixture samples were extracted at different locations in th e recirculation region and the shear layer formed behind a rearward-facing step in supersonic flow, and analyzed by mass spectrometry to determine the species co ncentration distributi on in the region. The point-wise mass spectrometer measurements we re complemented by acetone planar laserinduced fluorescence (PLIF) measurements to get a planar distri bution of fuel mole fraction in the recirculation region. Non-reacting flow test s and combustion experiments were performed by varying various fuel rela ted parameters such as injection location, injection pressure and fuel type. Fuel injecti on upstream of the step was not effective in

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x supplying enough fuel to the recirculation region and did not sustain the flame in combustion experiments. Fuel injection at the step base was effective in sustaining the flame. For base injection, the local fuel mole fraction in the recirculation region determined from experiments was an order of magnitude higher than the global fuel mole fraction based on total moles of air flowing through the test sect ion and total fuel injected in the test section. This suggests substant ial difference in flame stability curve for nonpremixed conditions in the scramjet engine compared to premixed flow. For base injection, fuel remained in the recirculati on region even at higher injection pressure. Due to slower diffusion rate, the heavier fuel had higher local mole fraction in the recirculation region compared to lighter fuel for a unit global fuel mole fraction injected in the test section. Hence fuel molecular we ight will affect the non-premixed flame stability limits in scramjet engine; the heavier fuel will have better fuel-lean and worse fuel-rich stability limit compared to lighter fu el. This is in addition to the fact that a lighter fuel such as hydrogen has a much wide r flame stability limit than a heavier fuel such as propane. The data obtained in the study can help develop a stability parameter for non-premixed flames and valid ate computational models.

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1 CHAPTER 1 INTRODUCTION Hypersonic Air-Breathing Vehicle Hypersonic air-breathing vehicles for laun ching payloads in space are an area of active interest in aerospace communities around the worl d. Such a vehicle could be used as the first stage of a two-st age-to-orbit launch vehicle with rocket-powered second stage required to reach orbit. In an optimistic scenario, the hypersonic air-breathing vehicle could also be used as a single-stage-to-orbit launcher. The existing rocket-based launch vehicles [Figure 1-1 (a)] carry fuel an d oxidizer along with them, and the latter contributes substantially to th e launch vehicle total weight. Th e oxidizer weight can be as high as 65% of the total weight while the usef ul payload is only about 2-3% of the total weight. This results in high launch costs and limits frequent access to space. An airbreathing launch vehicle [Figure 1-1 (b)] that propels itself by using oxygen from the atmosphere instead of carrying it onboard would free up a substantial portion of its weight that could potentially be utilized to significantly increase the useful payload weight as a fraction of total vehicl e weight and reduce launch costs. However, there are several challenges in realizing such a hype rsonic air-breathing vehicle. It will fly in the atmosphere at hi gh Mach numbers for a much longer time than a rocket, and thus will encount er significant drag heating on its airframe. Hence the airframe should be made of materials that can withstand extremely high temperatures; it also needs to be actively c ooled by cryogenic fuel circula ting under the vehicle skin. The air-breathing engine required to propel the vehicle at such high speeds is a significant

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2 challenge in itself. It is instructive to not e that the maximum speed attained by an airbreathing engine powered aircra ft is Mach 3.2; hence no airbreathing engine exists today for propelling an aircraft at hypersonic speeds. Air-breathing Engine Classification Propulsion for an air-breathing launch ve hicle would involve combined-cycle engines that utilize different forms of ai r-breathing engines most suited for different stages of the flight envelope. The vehicle ma y take off from a runway like a conventional airplane using a turbo-machinery based air-br eathing engine. As the flight Mach number increases in the supersonic regime and the operating limit for turbine-based engine is reached, the engine shifts to a ramjet mode of operation. At even higher flight Mach numbers approaching hypersonic speeds, the engine operation sh ifts to a scramjet mode. Turbojet Engine A schematic diagram of turbojet engine is shown in Figure 1-2 (a). Subsonic air entering the engine inlet is slowed down to low subsonic speed by a diffuser. Its stagnation pressure is increased by a compressor. Heat is added to the high pressure air by burning fuel in the combustor and its stagnatio n temperature is raised further. Fuel and air are mixed before ignition for uniform heat addition to the airflow; the flame in the combustor is anchored by a bluff body flameh older. The high pressu re, high temperature air is expanded through the turbine; the wo rk done on the turbine is used to drive the compressor. The air exiting the turbine still ha s high temperature; this thermal energy is converted to kinetic energy by accelerating th e air through a converging nozzle. The net increase in momentum of ai r passing through the engine pr oduces the thrust needed to propel the aircraft.

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3 A turbojet engine can also be used to power a supersonic aircraft by a suitably modified inlet. Oblique shocks at the inle t along with the diffuser slow down supersonic airflow to subsonic speed suitab le for turbojet engine opera tion. As the airflow speed at the inlet increases and more compression is ac hieved by the shock waves, less pressure rise is achieved at the compressor. Slowi ng down increasingly higher speed air at the inlet converts more of kinetic energy to thermal energy and ma kes it hotter. Increasing its temperature further by adding heat in the co mbustor causes air to approach the heat tolerance limit of the downstream turbine bl ades. Hence as supersonic speed of the aircraft increases to about Mach 2.5, the turboj et engine approaches its operational limit. A ramjet engine is suitable for higher Mach numbers. Ramjet Engine A ramjet engine operates only when the vehicle has been accelerated up to supersonic speed. It has a simple configura tion and differs significantly from a turbinebased engine in the fact that it has no m oving parts. A schematic diagram of ramjet engine is shown in Figure 1-2 (b). A series of oblique shocks at the inlet followed by a normal shock at the diffuser throat slows the air down to subsonic speed. Sufficient pressure rise is achieved by the shocks such that the need for a compressor to raise the airflow pressure and an associated turbine to drive the compressor is eliminated. Heat is added to the air flowing through the combus tor at subsonic speed. The high temperature achieved due to slowing down of high-speed ai r helps combustion progress faster in the engine. A converging-diverging nozzle accelerates the increased enthalpy flow exiting the combustor and exhausts it out of the engine at supersonic Mach number.

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4 Scramjet Engine As the flight Mach number increases fu rther in the supersonic regime to about Mach 5, the operation of a ramjet engine beco mes increasingly inefficient. The stagnation enthalpy of air, comprising kinetic and ther mal energy, increases with increasing Mach number. Shocking such a high energy flow dow n to subsonic speed converts the kinetic energy into thermal energy and substantially rais es its static temperature. At such high temperatures, heat addition would only dissoci ate products into ions and would not add enthalpy to the flow. Also, a normal shock cause s a significant stagnation pressure loss at high Mach number that is not desirable. Due to the above mentioned constraint s, the high enthalpy airflow cannot be slowed down significantly by the engine inlet an d diffuser, and it enters the combustor at supersonic speed. So heat is added to the airflow as it flows through the combustor at supersonic speed, and such an engine is called a supersonic combustion ramjet or scramjet engine. A schematic diagram of the scramjet engine is shown in Figure 1-2 (c). The hypersonic Mach number airflow is slow ed down to superson ic speed by oblique shocks at the inlet and then by the diffuser. An isolator is placed before the combustor to prevent interactions propagating upstream from the combustor to the inlet. In the combustor, heat is added to air flowing at supersonic Mach number. Typically, the flow speed entering the combustion chamber of a sc ramjet engine is about 1/2 1/3 of the vehicle flight Mach number. The high enth alpy flow exiting the combustor is then accelerated out of the engine by the converging-diverging nozzle. There are many challenges in achieving supersonic combustion in a scramjet engine1, 2, 3, 4, 5. Heat needs to be added to supersonic flow in a stable and efficient manner without causing significan t stagnation pressure losses in th e process. Efficient fuel-air

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5 mixing and rapid heat release is desirable so that a short co mbustor length with favorable thrust to drag ratio can be realized. However, the high-spee d air has an extremely short residence time in the combustor, of the order of a few ms Hence the fuel gets a very short time to mix efficiently with air, ignite, unde rgo complete combustion and add enthalpy to the incoming flow. Since the airflow entering the combustor is supersonic, the static pressure and temperature in the combustion chamber is quite low and unfavorable for rapid chemical reaction. The high-speed airf low is compressible, which has an adverse effect on the mixing process. Stagnation pressu re loss due to heat addition to supersonic airflow is unavoidable. If the airflow in the combus tor is slowed down to Mach 1 due to heat addition, it results in thermal choking of the engine. In th at case, the disturbanc es in the combustor propagate upstream and affect the air intake at the inlet. The static pressure in the combustor increases significantly and can even lead to blowing up of the engine. Hence particular care should be ta ken to avoid thermal choking. This can be achieved by having a diverging cross-section in downstream part of the combustor, and by limiting the fuel supply to an acceptable threshold. The scramjet engine has a lower pressure rise of air entering the inlet, less efficient heat addition in the combusto r, and higher stagnation pressu re losses compared to a turbojet engine. Hence it has a lower thrust to weight ratio than turbojet engine. To overcome this limitation, bigger scramjet e ngines are required on a hypersonic aircraft. This increases the total weight of the aircraft. Flameholding in Supersonic Flow Flameholding is an important area of concern in a scramjet engine. A parameter relevant to flameholding is the Damkohler number ( Da ) and is defined as the ratio of

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6 flow residence time in the combustor and the fuel reaction time; the reaction time is the time it takes for fuel-air mixture to mix, ign ite and release heat by chemical reaction. reaction residenceDa (1-1) Flameholding is possible only when the residence time available is more than the fuel reaction time (Da > 1). As mentioned earlier, the flow has a very short residence time in the combustion chamber and is of the order of only a few ms. In comparison, the chemical reaction time scales for hydrogen ar e similar to the scramjet combustor flow residence time scales, while the chemical reaction time scales for hydrocarbons are much higher; hence not enough time is available for flameholding in a supersonic combustor. The short residence time of the flow needs to be increased; hence a solution is to create a slower, subsonic recirculation region in part of the flow. Flow speed in this region is favorable to anchoring the flame, and the reci rculation flow in this region serves to mix fuel and air together. The static temperature ri se in the recirculatio n region due to slowing down of high enthalpy flow reduces the chemical reaction time scale. Hence Da > 1 in the flameholding recirculation region. The flam eholding region thus formed serves as a reservoir of hot pool of radicals that sustains the flame in the scramjet combustor, and also as a supplier of radical s helping to propagate combustion in the main supersonic flow. The flameholding region discussed above can be created in various ways. In turbojet engines, a bluff body placed in the combustor main flow is used as a flameholder. Such an approach is not approp riate in a scramjet combustor since the bluff body would cause significant st agnation pressure losses in supersonic flow. Various flameholding geometries suitable for supersoni c flow are shown in Figure 1-3. One way

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7 of anchoring the flame in a scramjet engine is to create a recess in the combustor wall in the form of a rearward step [Figure 1-3 (a )] or a cavity [Figure 1-3 (b)]. A subsonic recirculation region forms behind the step or in the cavity and acts as the flame anchor. The main advantage of recess flameholders is that they do not phys ically obstruct the supersonic flow and hence avoid stagnation pr essure loss. However, they create a very hot region at the combustor wall that needs active cooling. Since recess flameholders are located at the wall, they may not able to extend the heat release deep into the main airflow thus resulting in heat addition only to part of the ai rflow. Another flameholder is a slender strut placed in the main flow [Figur e 1-3 (c)] and with an appropriate geometry designed to minimize stagnation pressure losse s. The flow separates behind the strut and forms a recirculation region at the base th at anchors the flame. The oblique shocks generated by the presence of the strut in supers onic flow raise the flow static pressure and temperature thus assisting in flameholding. However, a drawback is that the strut experiences very high temperatures sin ce it physically obstructs the high enthalpy supersonic flow, and hence n eeds to be actively cooled. All the geometries discussed above ha ve a common flameholding mechanism, which is described for a rearward step in Figure 1-4. The approaching boundary layer of the main airflow separates at the step and forms a shear layer be tween the supersonic flow and the subsonic recircul ation region. The shear layer is pushed towards the wall due to the supersonic flow expansion at the st ep base. An oblique shock is formed at the shear layer reattachment point. The mass flow rate of air supplied from the main flow into the recirculation region is governed by the shear layer, as is the supply of hot combustion radicals from the recirculation regi on out into the main flow. Fuel is usually

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8 injected in the recirculation region where it mixes with the ambient air through the shear layer formed at the fuel jet. The flow pattern behind the step is complex in nature with smaller recirculation zones formed at the co rners besides the primary recirculation region. Further, the figure depicts on ly the 2-D flow pattern; sec ondary recirculation regions exist at the side walls and the flow pattern is 3-D in nature. Hence combustion occurring in the flameholding region is non-premixed in nature. Cavity Flameholder Ben-Yakar and Hanson6 reviewed the flow field char acteristics of supersonic flow over a cavity. Based on the L / D ratio of the cavity, the flow can be characterized as open or closed. Flow over cavity with L / D < 7-10 is characterized as open flow; the free shear layer separated at the cavity front wall re attaches at the back wall. For small aspect ratio cavity with L / D < 2-3, transverse acous tic waves oscillate along the cavity depth. For larger aspect ratio cavity, the waves os cillate along the cavity length. Cavity flow with L / D > 10-13 is characterized as closed flow; the free shear layer reattaches at the cavity floor. Closed flow cavity experiences much higher drag than open flow cavity, the drag force increases with increasing L / D ratio. The flow residence time in a cavity increases with cavity depth and the mass entrai nment rate increases with cavity length. The longitudinal oscillations in the cavity are caused due to shear layer impingement on the back wall. This unsteadines s in flow field is attractive for promoting fuel-air mixing; however it is undesirable for flameholdi ng. The oscillations can be passively controlled by angled back wall that prevents reflected acoustic waves. Active flow control over the cavity can be achieved by upstream injection to enhance the shear layer growth rate and alter its in stability characteristics.

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9 Gruber et al.7 experimentally and computati onally evaluated the flow field properties of various cavity geometries in s upersonic flow. Reducing the back wall angle, defined with respect to horizontal, increased the drag coefficient and shortened the flow residence time within the cavity. The increase in drag is due to higher pressures acting over a larger fraction of back wall area. The reduction in residence time is explained by the structure of recirculating flow within the cavity. Simulations showed that a primary recirculation region and a secondary embedde d vortex exist within the cavity. As the back wall angle is reduced, the primary reci rculation region increases in size and the secondary recirculation zone is diminished. For such a cav ity, the mass exchange takes place between the primary recirculation zone and the high speed mainstream flow. For a cavity with higher back wall angle, the mass exchange is slower since part of the exchange takes place between the low speed recirculation regions. Hence the residence time reduces with reducing the cavity back wall angle due to higher mass exchange rate. Ali and Kurian8 used fuel injection into a cavity in supersonic flow as an active control mechanism for enhancing the air entrainment rate into the cavity due to interaction between the fuel je t and free shear layer. Fuel in jected at different locations from the cavity floor and front/back wall incr eased the cavity pressure for all cases. The pressure fluctuations were suppressed for so me of the fuel injection locations. Mathur et al.9 performed supersonic combustion experiments with cavity flameholder; fuel was injected upstream of the cavity and from the cavity floor. The freestream conditions were varied to simulate different stages of hypersonic flight, and fuel was injected over a range of equivalence ratios. Stable flame anchored in the cavity

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10 and extending out into the main airflow was observed for all experimental conditions. The engine thrust increased with increasing fuel equivalence ratio. Yu et al.10 performed supersonic combustion experiments with various cavity flameholder geometries. Cavity length, back wall angle were varied and cavities in tandem were used. All cases showed a substa ntial increase in combustor pressure and temperature in comparison to the baseline case without cavity. This suggests enhanced volumetric heat release in the combustor aide d by the presence of cavities. Some cavity configurations such as the inclined cavity and the one with two cavities in tandem showed a better performance in compar ison with other configurations. Yu et al.11 performed supersonic combus tion experiments with cavity flameholder; kerosene was injected upstr eam of the cavity a nd piloted by hydrogen. Barbotaged atomization of liquid kerose ne using hydrogen increased the combustion efficiency substantially compared to pure liq uid atomization. The combustion efficiency for fuel injection perpendicular to the airf low was higher than angled injection due to deeper penetration into the airflow; howev er it was achieved at the cost of higher stagnation pressure loss. Combustion perf ormance improved with increasing cavity depth, which increased the flow residence ti me, but did not vary much with further increase beyond a certain depth. Tandem cavitie s performed better than a single cavity. Owens et al.12 performed supersonic combusti on experiments to examine the flame stability characteristics of cavity flam eholder. Kerosene was injected upstream in the flow boundary layer and piloted by hydrog en injected in the cavity. Temperature measurements in the recirculation region indi cated that cavities were regions of rich mixtures and their flame stabil ity was strongly affected by ai r stagnation temperature via

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11 changes in the local equivalence ratio. At high air stagnation temperature, rich mixtures existed in the cavity and injection of kerosene resulted in flame extinction, except for large hydrogen flow rates that enlarged the recirculation region a nd entrained additional air from the main flow. At lo w air stagnation temperature, the flame held even for large kerosene flow rates. Davis and Bowersox13 accessed flameholding properties of cavity in supersonic flow using a simplified, perfec tly stirred reactor analysis. Self-ignition of hydrogen was achieved at a lower temperature than hydro carbons; hydrogen had shorter ignition delay time than hydrocarbons at the same temperatur e. Heat loss from the cavity reduced the flammability limits. Once the lower residence time tr is calculated using the model, the cavity depth D required for flameholding can be estimated as D = tr U / 40, where U is the main airflow velocity. Strut Flameholder Brandstetter et al.14 established flameholding in supe rsonic flow in the recirculation region between a strut and a cylinder placed downstream of the strut. The cylinder surface temperature needed to be above a th reshold temperature for flameholding to be sustained after the ignitio n source was switched off. Northam et al.15 evaluated the performance of various strut geometries in supersonic flow. The strut incorporating the deep est step with perpendicular fuel injection downstream of the step exhibited the best co mbustion efficiency. The strut configuration with staged perpendicular fuel injecti on exhibited the lowest auto-ignition and flameholding limit. Hence an optimum strut co nfiguration should incorporate a deep step with staged perpendicular fuel injection fo r best mixing and flameholding performance.

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12 Lyubar et al.16 decelerated a small fraction of supersonic airflow in a scramjet combustor to subsonic conditions to provide a stable ignition and flameholding zone. This was achieved by an injector with its in ner surface shaped like a supersonic diffuser, combustor and nozzle. Hydrogen was injected pe rpendicular to the decelerated airflow in the injector and the main airflow. The strong temperature rise caused due to deceleration of the airflow in the injector resulted in auto -ignition and a sustained s upply of radicals to the main airflow. The normal shock causing au to-ignition in the injector was prevented from moving upstream during combustion by pr oviding an orifice to allow for pressure release from inside the injector to the main airflow. The mass flow from the orifice also enhanced fuel-air mixing in the main airflow. Gerlinger and Bruggermann17 performed numerical simulations for strut flameholder in supersonic flow with hydrogen injected in the fl ow direction at the base of the strut; the effect of varying the lip thickn ess at the injector end was investigated. The mixing efficiency was nearly independent of lip thickness variation. Changes caused within the shock wave/expansion pattern at the injector exit due to lip thickness variation had a moderate influence on the stagnation pressure loss. Tabejamaat et al.18 performed numerical simula tions and experiments to investigate the effect of geom etry variation of a strut in supersonic flow. Hydrogen was injected parallel to the airflow and into the recirculation region formed at the base of the strut. Increasing the base height increased the recirculation region size, improved the mixing efficiency and hence increased th e maximum combustion temperature in the recirculation region. Increasing the slit width of injector exit increased the recirculation region size, but did not appreciably aff ect the mixing efficiency and maximum

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13 combustion temperature in the recirculation region. The double slit injection changed the flow field compared to single slit injection; increasing the distance between the two slits beyond a certain limit changed the flow field s ubstantially and caused flame extinction. Gruenig et al.19, 20 performed experiments to investigate flame stabilization and mixing by pylons in supersonic flow. Pylons can be considered as short-length struts. Mixing efficiency and combustion performa nce of various pylon geometries were investigated. The best com bustion performance was exhibi ted by the pylon with fuel injected inclined to the airflow and through ramps designed to create stream-wise vortices that promote mixing. Flame stabilization was achieved by oblique shocks generated by the pylon; the shocks increased the flow te mperature and pressure and caused autoignition of fuel-air mixture after they had mixed downstream of the pylon. A wedge appropriately positioned downstream of the pyl on was used to modify the shock structure and induce auto-ignition at a shorter com bustor length. The shock train was in turn influenced and pushed upstream by heat release in the combustor. Oblique shocks are often formed within a scramjet combustor. Huh and Driscoll21 investigated the beneficial e ffects of shock waves on a supe rsonic jet flame. Shock waves optimally positioned with respect to the jet fl ame altered the flame size and significantly improved the blowout limits. The shock wave s enhance mixing by directing the airflow radially inwards towards the fuel and thus in crease the air entrainment rate. They create an adverse pressure gradient which increases the recirculation regi on size. They also improve the chemical reaction rates by increa sing the static temper ature and pressure. Free Shear Layer Fluid Dynamics The free shear layer plays a crucial role in bringing air from the main flow into the flameholding recirculation region and in tr ansporting combustion ra dicals out of the

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14 recirculation region into the main airflow. Hence the free shear layer fluid dynamics is discussed here. Dimotakis22 reviewed experimental data on turbulent free shear layer growth, mixing and chemical reactions formed between two uniform gas streams. The local Reynolds number is given by (1-2) where U1, U2 are the velocities of the two streams. The chemical product formation at a station x can be expressed as (1-3) where the first factor measures the shear layer growth, the sec ond shows mixing within the shear layer, and the third indicates the chemical products formed within the molecularly mixed layer. Shear layer growth: It depends on several flow para meters like velocity ratio (r), density ratio (s), convective Mach numbers of the two streams (Mc1, Mc2), relative mean density reduction attributed to he at release for combusting flow (q), and pressure gradient. (1-4) where Uc is the velocity of the large-scale turbulent structures in the shear layer. The incompressible shear layer growth rate increas es with an increase in density ratio. The incompressible shear layer growth rate as a function of velocity and density ratio is expressed as follows. ) / ( ) / ( ) / ( /m P m Px x 2 1 410 / Re U U U U 0 0 2 2 2 1 1 1 1 2 1 2/ ) ( / ) ( / ) ( / / q a U U M a U U M s U U rc c c c

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15 1/21/21/2 1/2(1)(1)(1)/(1) (,)[1] 2(1)12.9(1)/(1) rsss rsC x srrr (1-5) Compressibility has an adverse effect on shear layer growth. For compressible flow, the growth rate decreases with increasing convec tive Mach number as s hown in the relation below. 2 0 ) 1 ( ) ( ] 0 ; [ ] ; [2 13 1 1 1 f f e f M f M s r x M s r xcM c c c (1-6) Heat release in subsonic flow results in a moderate decrease in shear layer growth. The outward displacement veloci ty due to heat release impe des the entrainment process and offsets the dilatation effect. The density reduction approaches a limiting value with higher heat release due a substantially unmixed fluid at high Reynolds number. Pressure gradient results in an increase or decrease in shear layer growth depending on velocity and density ratio. Mixing: Enhanced mixing of the two free streams is obtained when shear layer instability causes a transition from large, tw o-dimensional vortices to three dimensional, fully developed, turbulent flow in the shear la yer. However, at high Reynolds numbers, a substantial fraction of fluid in the sh ear layer is not molecularly mixed. The shear layer entrains fluid from the tw o free streams in an asymmetric way. The molar entrainment ratio En for compressible shear layer is expressed as follows. 68 0 ) 1 )]( ( 1 1 1 [2 / 1 1 1 2 l c l nC s M f r r C M M E (1-7) where M1, M2 are molecular weights of the two streams.

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16 The probability density function (PDF) of concentration measurements across the width of a shear layer shows that away from the boundaries of the two free streams, the high-speed fluid fraction distribution in the mixed shear layer tends to a most probable value of E defined as follows. (1-8) Pitz and Daily23 made 3-D velocity measurements in the free shear layer formed behind a rearward step in subsonic flow. The shear layer growth rate was higher compared to other measurements for parallel streams; this was attributed to reverse flow in the recirculation region behind the step. Th e shear layer growth rate was same for nonreacting and reacting flow. Flame Stability Limits Flame Stability in High-Speed Subsonic Flow Ozawa24 compiled experimental results on wake stabilized flames in high-speed combustion systems. The limitation of the compile d results is that they are applicable to premixed, subsonic flows while the combustion process inside a scra mjet engine is nonpremixed and supersonic. A bluff body placed in a combustion chamber stabilizes the flame in its wake. They are axisymmetric in shape (cone, hemisphere, disc) and also two-di mensional (V-gutter, cylinder, flat plate) with varying degree of bluffness. Flame stabilization depends on aerodynamics of the flow in the wake of the flameholder. The mass entrainment rate from the main flow into the recirculation region formed behind the body increases and the flow residence time in the region decreases with an increase in the degree of bluffness. However, a compromise has to be made be tween improved flameholding characteristics 1 n n EE E

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17 and higher pressure losses of a more bluff body. A 2-D body has a higher mass entrainment rate in its wa ke and a longer residence time than a corresponding axisymmetric body, and hence has a better flameholding performance. Combustion reduces the mass exchange rate and increase s the residence time in the recirculation region. Various parameters of the inlet flow aff ect the flame blowout velocity: pressure, temperature, turbulence, water vapor contamination. An increase in inlet flow pressure and temperature has a favorable effect on the blowout velocity, with temperature having a stronger influence. Higher free-stream tur bulence intensity increases the mass exchange rate behind the flameholder and the resulti ng quenching effect decreases the blowout velocity. Higher water vapor content of the in let flow increases the ignition delay time of the fuel and hence decreases the blowout velo city. A stability parameter was formulated consisting of inlet flow velocity, pressure, temperature, and the flameholder type. The flameholder parameter is the maximum width of the recirculation region formed behind it. A plot of equivalence ratio vs. stability parameter gives the flame stability curve for a given fuel. The correlation be tween stability parameter ( SP ) and equivalence ratio is shown below. ) ( ) )( )( 1 ( ) 300 (0 5 1 0f d d U d atm P K T SP (1-9) where : premixed equivalence ratioT0 : free stream stagnation temperature P : free stream static pressure U : free stream velocity d0 / d : flameholder shape parameter

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18 The mixing intensity in the recirculation region behind 2-D flameholders is about 30% higher than axisymmetric flameholders However, the highest mixing intensity achieved in stirred reactors is two to three times higher th an bluff body flameholders. Huellmantel et al.25 performed experiments for st abilizing premixed, subsonic, propane-air flames using a cavity recess in the combustor wall as the flameholder. Flameholding performances of various cavity geometries were studied. The flame stabilization curve was generated for each cav ity by obtaining the flow velocity at which flame blow out occurred as a function of both fuel-lean and fuel-rich equivalence ratios. A long cavity had a wider flame stability ra nge than a short cavity and a deep cavity performed better than a shallow cavity. It indicates the necessity of having sufficient recirculation volume for achieving good stability and s hows that increasing the size of the flameholder increases its performance. The de gree of slope of the downstream end of the cavity did not affect flame stabilization appreciably. The cavity flameholder had a superior bl owout performance when compared to a 900 V-gutter, which is a standard bluff body flameholder. Also, the cavity caused much less pressure loss in the main flow since it did not physically obstruct the flow. However, a cause for concern in using a cavity flameholder is the excessive heating of engine wall in its vicinity. Baxter and Lefebvre26 determined the fuel-lean flame blowout limits for high-speed subsonic afterburner combustor systems by vary ing a range of parameters such as the Vshaped flameholder geometry, injector-fla meholder spacing, airflow Mach number and stagnation temperature. The effects of cha nging some of these parameters were at variance with homogenous results27. The heterogeneous flame blowout equivalence ratios

PAGE 29

19 were leaner than the homogenous values due to richer local fuel -concentration in the flameholding region. An approximate analys is was developed to model heterogeneous effects such as transport and vaporization of liquid fuel droplet s from the injection location to the flameholder, droplet capture and vaporization at the flameholder surface, and gas entrainment into the recirculation re gion. The effective equivalence ratio in the recirculation region determined from the above analysis and the lean blowout equivalence ratio for homogenous flame obtai ned from correlation of experimental data27 were in good agreement with each other. Non-Premixed Flame Stability in Supersonic Flow Niioka et al.28 performed flame stabilization tests in supersonic flow using a strut divided stream-wise into two parts, with hydr ogen injected in the spacing between them. For different spacing distances between the strut components ( L ), flame stability plots were generated by determining the blowout fu el flow rates as a function of airflow stagnation temperature. The flameholding perf ormance of the strut system was found to depend greatly on the distance L The shock/expansion waves formed around to the struts were not observed to vary much with variation in L Hence the variation in flameholding performance was due to the competition be tween flow residence time and chemical reaction time in the recirculation region formed in the intermediate region between the parts. For short and long distances L the recirculation region was too fuel rich and fuel lean respectively resulting in large reaction times ( Da = residence / reaction < 1), and hence did not give good flame stabil ity. For moderate distances L air entrainment from supersonic airflow into the recirculation re gion was adequate and reasonably well mixed.

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20 The reaction time was less than residence time ( Da > 1), hence flame stability was achieved. Zakkay et al.29 measured the residence time of a gas introduced in the recirculation region behind an axisymmetric body in s upersonic flow for various experimental conditions. The mo lar concentration of the gas in the recirculation region decayed exponentially with time. The dissipation of gas from the recirculation region was due to diffusion process. The residence time in the recirculation re gion was of the order of a few ms and t U / D ~ 75 for laminar and 40 for turbulent flow, where t residence time in the recirculation region, U freestream velocity, D bluff body base diameter. The residence time is lower for turbulent flow than laminar flow due to higher diffusion rate for turbulent flow compared to lamina r flow. The residence time was independent of the concentration of gas in the recirculation region; higher concentr ation gas di ssipated faster than lower concentration gas thus re sulting in the same residence time for both cases. Driscoll and Rasmussen30 performed an analysis to develop a correlation for nonpremixed flame stability limits in supersonic flow. The analysis is based on the idea that the flame base is sustained in the shear laye r and not in the recirc ulation region. Flame propagation speed along the stoi chiometric contour in the shear layer is matched by the velocity of the incoming gas. Hot products in the recirculation zone preheat the shear layer gases and increase the flame propagation speed. Flame blowout is governed by the imbalance between flame propagation spee d and gas velocity. Some additional parameters governing blowout appeared in th e non-premixed flame correlation compared to the premixed flame correlation: fuel inje ction location relative to the recirculation

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21 region, fuel injection temperat ure, and stoichiometric fuel mixture fraction. Also, a global equivalence ratio appeared in the correlation th at is different from the local equivalence ratio in the flameholding recirc ulation region. The correlation was applied to previously measured data covering supersonic and subsoni c flows; cavities, bluff bodies and struts as flameholders; hydrogen, ethylene, methane and propane as fuels; and had an average uncertainty of 55%. Rasmussen et al.31 examined the stabili ty of hydrocarbon flames in supersonic flow using cavity flameholders. The effect of the different parameters on lean and rich flame blowout limits was investigated: fuel type (ethylene, methane), cavity geometry (rectangular, aft wall ramp cavity), fuel injec tion location (cavity fl oor, cavity aft wall), and airflow Mach number. The blowout lim its showed strong dependence on fuel injection location. For lean blowout limit, the aft wall ramp injection performed better than floor injection. This is because ramp injection puts fuel directly into the main recirculation region formed in the cavity, and hence gets dist ributed throughout the cavity. Whereas floor injection puts fuel only in the shear layer, henc e fuel is unable to reach much of the cavity. For rich blowout lim it, the floor injection performed better than aft wall ramp injection. The reason is the sa me as explained above. For floor injection much of fuel bypasses the cavity and remains concentr ated in the shear layer, while ramp injection floods the cavity with fuel by inject ing directly into the recirculation region. Hence the fuel injection location relative to the recirculation region is an important parameter in determining flame stability lim its. Ethylene had better flame stability than methane for lean as well as rich limits, sin ce it has a much shorter ignition delay time and faster flame propagation speed. The airflow Mach number had little effect on lean

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22 stability limit, however Mach 2 flow showed better flame stability than Mach 3 flow for rich limit. Increase in Mach number reduces the flow pressure, temperature and time scale, and its effect on flame stability can be better understood by computer simulation. Winterfeld32 performed flame blowout experime nts in supersonic flow using a contoured cylindrical flameholder and a c one-cylinder flameholder. A recirculation region formed behind the flameholder. Hydrogen was injected into the recirculation zone at different angles re lative to the airflow ( ). The measured recirculation region size was bigger for = 00 fuel injection angle than for = 900. This is because more fuel was injected directly into the recirculation region for = 00 while part of the fuel was blown into the supersonic airflow for = 900. The flame blowout curve plotting normalized fuel flow rate as a function of blowout airflow Mach number was also sensitive to the fuel injection angle. For fuel rich limit, better flame stability was achieved for higher injection angles. It is again explained by the f act that fuel injection at low angle floods the recirculation with fuel, while at high angle part of fuel escapes into the supersonic airflow. Rasmussen et al.33 investigated the effect of fuel injection pressure on flame location within a cavity in supersonic flow. Fuel was injected from the cavity floor and from the cavity aft wall; the flame chemiluminescence was captured using a digital camera. For moderate fuel flow rates with both injection locations, the flame was anchored in the shear layer and also extended to the recirculation region. For lean and rich blowout with floor injection, the flame structure was similar to the moderate fueling case. For lean and rich blowout with aft wa ll injection, the flame structure showed a marked departure from moderate fueling case ; the flame was not anchored in the shear

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23 layer and was restricted to a region close to the aft wall. The change in flame stabilization mechanism for aft injection may be attributed to a change in the cavity flow field with heat release. Species Distribution in Non-Premixed Fl ameholding Region in Supersonic Flow Experiments Hsu et al.34 used Raman scattering to make fuel distribution measur ements inside a cavity in Mach 2 non-reacting flow. Ethylene wa s injected at a low angle upstream of the cavity. The effect of fuel injection pressure cavity size, and imposed back-pressure on fuel transport in the cavity was studied. Fuel-rich pockets were observed in the cavity. For upstream fuel injection, mass transport through the shear layer and its interaction with the cavity aft wall cont rols the amount of fuel enteri ng into the cavity. The amount of fuel entering the cavity decreased as the fu el pressure was increased from moderate to high values. Higher fuel injection pressure w ith an increased jet momentum penetrates into the main airflow and less fuel gets entrained into the shear layer and reaches the cavity. An increase in cavity size captured more fuel at the ca vity back wall, but the drag penalty also increased. Increase in back pr essure, which simulates combustion conditions, caused the boundary layer upstream of the cavity to separate thus changing the shear layer interaction with the cavity. The fuel jet pe netrated into the main flow due to reduced momentum of the boundary layer an d less of it reached the cavity. As a follow up of the experiment s described above, Gruber et al.35 examined the effect of fuel injection location, fuel flow rate, and induced back -pressure on the cavity flameholder performance in supersonic flow NO-PLIF was used to visualize fuel distribution in the cavity fo r non-reacting flow, and OH-PLIF was used in combustion experiments. From non-reacting flow PLIF vi sualizations, it was concluded that upstream

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24 fuel injection with passive entrainment of fu el from main airflow into the cavity is less desirable than direct fuel injection into the cavity. Unlike upstream fuel injection, direct injection from the cavity aft ramp into the main recirculation region provided a spatially uniform fuel-air mixture in the cavity. Al so, the cavity fuel distribution for direct injection remained relatively insensitive to changes in the main airflow as simulated by induced back-pressure. In combustion experiment s, as the fuel flow rate from cavity aft ramp was increased, the cavity was flooded with fuel and adversely affected combustion in the region. However, when a shock train was imposed for the high fuel flow rate, it significantly improved cavity combustion by caus ing the cavity shear layer to separate which effectively increased air entrainment in the cavity. The shock train could also enhance fuel-air mixing in the cavity. Uchiumi et al.36 followed up on the investigation by Niioka at al.28 discussed earlier and conducted experiments to improve the fl ameholding performance of a strut divided into two parts in supersonic flow. Previous non-reacting flow measurements of local equivalence ratio along the distance between th e strut parts showed th at for short distance L the intervening region was largely fuel ri ch in composition. Such a mixture required a long chemical reaction time compared to th e short residence time available, and hence flameholding could not be es tablished. As the distance L was increased, the measured local equivalence ratios were reduced since mo re air from the main flow was entrained into the recirculation region. However, fo r a certain range of moderate distance L the chemical reaction time was more than the fl ow residence time a nd hence flameholding could not be established. M odifications in the fuel-injecting strut were made in the present experiment. For short distance L between the struts, slits incorporated in the

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25 injection strut brought air from main flow into the recirculat ion region and diluted its fuel rich composition. For moderate distance L a recess was provided around the fuel injector to prevent excess air from entering the intervening space. The strut modifications resulted in Da > 1 and hence flameholding was established in both cases. Zamma et al.37 measured pressure and gas composition for non-reacting supersonic flow over a step with fuel injected downstream of the step and transverse to the airflow. Species in the recirculation re gion behind the step were extract ed at the wall and analyzed by gas chromatography. Fuel wa s entrained from the jet into the recirculation region behind the step. The fuel volume fraction in the recirculation region decreased with increasing fuel-air dynamic pressure ratio. Hi gher fuel volume fraction was observed in the recirculation region for li ghter fuel compared to heavier fuel injected at the same dynamic pressure ratio. The gas composition measurements were used to estimate the ignition characteristics of the flameholder38, 39. Thayer and Corlett40 measured pressure, temperature and gas composition in the separated recirculation region upstream of a fu el jet injected transverse to non-reacting supersonic airflow. The species extracted at the wall were analyzed by gas chromatography. A large part of the recirc ulation region had a nearly uniform fuel distribution; the region was fuel-rich in composition. The fuel concentration in the recirculation region decreased with decreasing fuel mass flow rate. It was estimated that approximately 5 % of the injected fuel was en trained in the recirculation region. McDaniel et al.41 made 3-D measurements of flow variables for a flow field with staged transverse fuel injection downstream of a step in non-reacting supersonic flow. Pressure, temperature, velocity and fuel mo le fraction were measured using LIF. Fuel

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26 was entrained into the recirculat ion region behind the step from the jet closer to the step base; fuel concentration close to stoichiometr y was observed in the recirculation region. Strokin and Grachev42 obtained experimental data on ignition and flameholding in supersonic flow using a cavity flameholder fo r a variety of experi mental conditions; the results were also reported by Ogorodnikov et al.43. The data was correlated to obtain a flame stability curve using the air/fuel equiva lence ratio in the cavit y recirculation zone and a flameholding parameter based on the airflow velocity, airflow stagnation temperature, recirculation region static pre ssure and cavity length. The equivalence ratio in the recirculation region was empirically estimated using air and fuel stagnation temperature, recirculation re gion static temperature, ap proaching airflow boundary layer thickness, spacing between fuel in jection holes, and overall air/fu el equivalence ratio. Morrison et al.44 performed experimental and an alytical studies on a dual-mode propane-fueled ramjet/scramjet combustor to determine the conditions necessary to establish flameholding in the engine. The ef fects of non-premixed fuel-air mixture and combustion-induced shock train on flame holding were investigated. Hydrogen was injected near the exit to raise the back-pressure and establish a shock train that resulted in subsonic flow over the step flameholder a nd simulated ramjet conditions. Propane was injected at a low velocity at the step base to prevent it from penetrating through the recirculation region. The air entrainment rate into the recirculation region was estimated analytically from the step base pressure measurements. Premixed conditions were assumed to exist in the recirculation region a nd the local equivalence ratio in the region was calculated from the known fuel flow rate and estimated air entrainment rate. Lean blowout at a local equivalence ratio close to unity indicated poor fuel-air mixing in the

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27 recirculation region; the mixing efficiency was estimated as 30%. The experiments were repeated for different levels of back-pressu re to determine both fuel-lean and fuel-rich blowout points and construct a flame stability curve. The curve was similar in shape to the Ozawa stability loop fo r premixed subsonic flow24, but it had reduced blowout limits. Computations Kim et al.45 reported 2-D numerical simulation of fuel injection over a cavity in non-reacting supersonic flow. Wh en fuel was injected upstream of the cavity, it drastically suppressed flow oscillations in the incoming boundary laye r leading to a more stable free shear layer over the cavity. For downstream injection, the interaction between flow oscillations near the aft cavity wall and th e injected jet amplified the cavity pressure oscillations as compared to the case w ithout injection. Fuel mass fraction contours showed that large vortices occurred that pulled the fuel upstream into the cavity. Correa and Warren46 performed 2-D numerical simulations and experiments for non-reacting supersonic flow over a backward-f acing step. Fuel was injected downstream of the step and transverse to the airflow. Th e computed fuel mixture fraction showed that the fuel was confined to the wake of the st ep. In reality, a 3-D horse shoe vortex forms around the fuel jet and it penetrates further into the main airflow. Glawe et al.47 performed 2-D numerical simulations for helium injected at the base of a strut in non-react ing supersonic flow. CFD cont ours showed a high helium mole fraction distribution in the re circulation region formed at the base of the strut. Mass Sampling for Species Concentratio n Measurement in Supersonic Flow Mitani et al.48 performed experiments and comput ations to analyze and select a suitable mass sampling probe for a hydrogen supersonic combustor. The extracted samples were analyzed for species compos ition using gas chromatography. Different

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28 types of gas sampling probes were examined: a reaction-oriented probe having a uniform cross-section with no water cool ing; a reaction-freezing orient ed water cooled probe with a short cross-section expanding to a larger cross-section to fa cilitate expansion cooling of the sampled gas; and a static pressure-type probe to grasp the external flow across the boundary layer on probe surface. The effect iveness of freezing-oriented probe was examined by comparison with reaction-orie nted probe. Combustion efficiency as a function of stagnation temperature T0 was estimated from gas sampling by each of the probes. The reaction-orie nted probe allowed reactions to co ntinue ahead of and inside it and showed complete combustion for T0 > 910 K. It gave high but false values for combustion efficiency. The freezing-orient ed probe indicated that combustion was initiated when T0 > 1200 K and gave lower combustion efficiency, which were in line with static-type probe measur ements. The validated gas sampling was applied to scramjet testing under flight Mach numbers up to 8. Th ey clarified distortions of air and hydrogen in the swept-back, side-compression type engines. Chinzei et al.49, Masuya et al.50, Ciezki et al.51, Rogers52 used mass sampling and subsequent analysis by gas chromatography to determine species distribution at different cross-sections in a supersonic combustor for non-reacting and reacting flow. The mixing effectiveness of various fuel injection conf igurations such as step injection, strut injection, and wall injection normal to the airflow were examined. Ng et al.53 developed an aspirating hot-film mass sampling probe and used it to measure local gas composition in a helium-a ir supersonic shear layer. The probe consisted of a diverging section followed by a constant area section and a choked orifice, and was connected to a vacuum pump at the back. The configurati on allowed the sampled

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29 flow to expand and form a normal shock in the diverging section of the probe, thus avoiding a normal shock and flow spillage at the probe tip. The hot-film sensor was placed in the constant area section of the probe and was connected to an anemometer. The voltage response of the anemometer is a function of the sampled gas composition, total pressure and total temperature experi enced by the sensor. The total pressure and temperature at the sensor were measured separately, which allowed for the determination of sampled gas composition using calibration curves. Cox et al.54 used the probe to study mixing efficiency of an aerodyna mic ramp in supersonic flow. Optical Diagnostics for Species Concen tration Measurement in Supersonic Flow Planar Laser-Induced Fluorescence (PLIF) Hanson et al.55, Schulz and Sick56 provided an overview of planar laser-induced fluorescence (PLIF) as a diagnostics tool for measuring species concentration, temperature, pressure and velocity in the flow field. The fluorescence signal Sf can be expressed as follows. Q A A kT P dV hc E Sabs c optics f / (1-10) where optics : efficiency of the collection optics E : laser fluence (J/cm2) hc / : energy of a photon dVc : collection volume (cm3) : mole fraction of absorbing molecule abs : molecular absorption cross section (cm2) A : spontaneous emission rate (s-1) Q : collisional quenching rate (s-1)

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30 The ratio A / (A + Q) gives the fluorescence yield which is the fraction of absorbed photons re-emitted as fluorescence photons. depends on composition, temperature and pressure of the gas mixture, and excitation wavelength. abs depends on gas temperature and excitation wavelength. Hence the fluoresce nce signal is a function of species mole fraction, temperature, pressure and excita tion wavelength. Mole fraction imaging is straightforward for a flow with relatively constant temperature and pressure. For a relatively constant pressure flow with temp erature variation, the mole fraction can be determined by selecting a suitable excitatio n wavelength that mini mizes the signals overall temperature dependence in the te mperature range of experiment. Another approach is to take two almost simultaneous images with excitation at two different wavelengths. The ratio of the two image signa ls is a function of temperature only. The flow field temperature variati on thus obtained can be used to determine the mole fraction distribution from either of the images. Fox et al.57 used NO PLIF to determine the fu el mole fraction distribution and evaluate the mixing performance of various fuel injectors in supersonic flow. The temperature dependence of the signal was minimized by tuning the laser to excitation wavelengths for which fluorescence is rela tively independent of temperature. The pressure dependence of the signal was mi nimized by canceling the implicit pressure dependence of fluorescence yield with explicit pressure dependence of number density; this was achieved by increas ing the quenching cross secti on to an order of magnitude higher than the fluorescence cross section.

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31 Hartfield et al.58, Abbitt et al.59, Hollo et al.60, Hartfield et al.61 used iodine PLIF to measure injectant mole fraction distributi on in non-reacting supersonic combustor and evaluate mixing rates for different injection configurations. The fluorescence intensity was a function of injectant concentration as well as the local thermodynamic pressure and temperature. The ratio of the fluorescence signal collected with only the injectant jets seeded to the signal collected with both the jets and the main flow seeded was taken to cancel the thermodynamic dependence of fluore scence, and the injectant mole fraction was obtained as follows. total f jet f total jet t injecS S C N N X tan (1-11) where X: mole fraction, N: number density, Sf : signal intensity. Lozano et al.62 explored the use of acetone (CH3-CO-CH3) as a suitable tracer for PLIF concentration measurements in gaseous flows. Acetone has a fairly high vapor pressure (180 torr) at room temperature ( 293 K) and hence allows a good seeding density, which can be increased further by increasing th e temperature. It absorbs over a broadband of wavelengths (225-320 nm) with a maxi mum between 270-280 nm (absorption crosssection = 4.7 x 10-20 cm2). The fluorescence emission is broadband in blue (350-550 nm) with peaks at 445 and 480 nm, quantum efficiency = 0.2 %, and a short lifetime of less than 4 ns. The fluorescence quantum yiel d is dominated by intersystem crossing; collision quenching is neglig ible and hence the yield is independent of local gas composition. For the experimental conditio ns investigated, the signal showed no dependence on temperature and he nce was only sensitive to the species concentration. Acetone phosphoresces in the bl ue, with a similar emission spectrum as fluorescence

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32 (350-600 nm) and a long decay time (200 s for 313 nm excitation). However, it is greatly quenched by trace amounts of oxygen and also depends on temperature. Bryant et al.63 conducted experiments to determine the temperature and pressure dependence of acetone LIF signal. The pressure and temperature range investigated were 0.1 1.0 atm and 240 300 K respectively, conditions akin to that in a supersonic wind tunnel. For laser excitation wavelength of 266 nm, the LI F signal showed no dependence on pressure and varied by 5% over the temperature range. Hen ce the signal can be used to directly obtain species concentr ation in an acetone-seeded flow. Thurber et al.64 investigated the temperature a nd excitation wavelength dependence of acetone LIF signal; the temperature range was 300 1000 K and the wavelength range was 248 320 nm. The signal vari ed with temperature and wavelength for the range investigated. At 266 nm, the signal per molecule was constant for 300 350 K and decreased for higher temperatures. VanLerberghe et al.65 used acetone PLIF to investigate mixing of an underexpanded sonic jet injected transversely in a supersonic crossflow. Instantaneous, mean, standard deviation images and image intens ity probability density functions (PDF) were used to study the mixing process. For the given flow conditions, PLIF signal intensity was proportional to the acetone mixture frac tion; temperature and pressure were estimated to not play a significant role. Hartfield et al.66 measured pressure, temperature and velocity in supersonic flow using iodine PLIF. The uncerta inty in measurements due to possible condensation of iodine in the low temperature supersonic ai rflow was investigated. Minor condensation was observed on the test section walls. The si gnal intensity from the condensed droplets

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33 would be much higher than that from vapo r. However, the measured signals were comparable with the calculated theoretical va lues. Hence no direct evidence of significant iodine condensation in the airflow was observed. Raman Scattering OByrne et al.67 used dual-pump coherent anti-S tokes Raman spectroscopy (CARS) to measure temperature and mo le fraction distribution of N2, O2 and H2 in a supersonic combustor. A parallel CFD study was done by Cutler et al.68. CARS has the advantage of producing a coherent signal beam in a particul ar direction. This increases the signal-tonoise ratio, and also allows measurements wh ere optical access to the flow is limited. Three lasers were used in the experime nts. The 532 nm beam output from a Nd:YAG laser was split three ways. The first beam was used as the green pump beam for N2 CARS. The second beam was used to pump a dy e laser operating at 554 nm that provided the yellow pump beam for O2 CARS. The third beam was used to pump another dye laser operating at 607 nm that provided the red stokes beam for both N2 and O2 CARS. The frequency difference between blue pump beam and red stokes beam equals a vibration Raman shift of N2, while the frequency difference between yellow pump beam and red stokes beam resonates with a vibration Raman shift of O2. The resulting blue CARS spectrum contains both N2 and O2 spectra. Coincidentally, several pure-rotational Raman transitions of H2 are also present in this spectral region. The relative intensities of N2, O2 and H2 spectra provided a measure of mole fractions of these species. Kasal et al.69 evaluated mixing of hydrogen injected from planar and lobed struts in supersonic non-reacting flow. 1D Raman spec troscopy was used to measure species composition at different cross-sections al ong the combustor. A 532 nm, 400mJ Nd:YAG

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34 laser was used as the exciting source. The resulting Raman spectrum was used to obtain mole fractions of H2, O2 and N2 in the flow. Scope of Study The aim of this study is to investigate the non-premixed conditions in the recirculation region formed behind a rearward st ep in supersonic flow; such a study is of practical importance in understa nding the flameholding mechanism in a scramjet engine combustor. The step is selected as the flame holder geometry since it is the simplest and most widely used configuration in a supersonic combustor. The literature review section discussed earlier shows that only a few inves tigations have been conducted to determine the local species distributi on in the flameholder recirc ulation region for non-reacting supersonic flow. As per the authors knowledg e, no such data has been reported for combustion tests. This study creates a database of local species mole fr action and fuel equivalence ratio distribution in the reci rculation region formed behi nd a step in supersonic flow. Such a database also helps in providing useful input data for complementary computational fluid dynamics (CFD) effort s. Both non-reacting and combustion cases were investigated. The airflow parameters su ch as Mach number, stagnation pressure and stagnation temperature were he ld constant. For a given flow condition, the influence of the following fuel related parameters on th e species distribution and combustion in the flameholding recirculation region were investigated: Fuel injection location relati ve to the recirculation region Fuel injection pressure Fuel type

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35 The tools used in the investigation were a combination of mass sampling and composition analysis using mass spectrometry (MS) and acetone pl anar laser induced fluorescence (PLIF) for visualizing species concentration distri bution in non-reacting flow.

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36 (a) (b) Figure 1-1. Comparison between (a) Rock et-powered vehicle [Source: www.isro.org] and (b) Scramjet-powered hypersoni c vehicle [Source: www.nasa.gov].

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37 Figure 1-2. Schematic diagram of (a) turboj et engine (b) ramjet engine (c) scramjet engine. Com p resso r Turbine N ozzle Inle t Combusto r M < 1 (a) M > 1 Combusto r Inle t N ozzle Oblique shock M < 1 (b) Oblique shock M > 1 Inle t Isolator Combustor M >1 N ozzle (c)

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38 Figure 1-3. Various flameholder geometries for supersonic flow: (a) rearward step (b) cavity (c) strut. Figure 1-4. Schematic diagram of su personic flow over a rearward step. (a) (b) (c) M > 1

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39 CHAPTER 2 EXPERIMENTAL SETUP Supersonic Wind Tunnel The supersonic wind tunnel facility used in the experiments, and shown in Figure 2-1 (a), provides direct conn ect tests with a variable combustion chamber entrance Mach number of 1.6 3.6 and sta gnation temperatures corresponding to Mach 3.0 4.8 flight. The flight Mach numbers correspond to the transition phase from ramjet to scramjet engine. The wind tunnel is a continuously operati ng facility using a vi tiated heater based on hydrogen combustion with oxygen replenishmen t, electronically controlled by a fuzzy logic controller70 to maintain a constant 0.21 oxygen mole fracti on at all conditions, and to maintain at the heater exit the constant stagnation temperature profile required for the experiment. A bell mouth with four-side cont raction leads to the supersonic nozzle with compression on two sides and interchangeable nozzle blocks that cover the range of Mach 1.6 3.6. All the experiments discu ssed here were performed with combustion chamber entrance Mach 1.6 and cold air (T0air = 300 K). A constant area isolator is placed between the nozzles and the combustor sec tion to protect the nozzle from upstream pressure rise due to combustion in the test section. Optical access is available to the isolators flow from three sides. The isolator cross-sec tion is 2.5 x 2.5 cm2 upstream of a rectangular, rearward facing step having step height H = 12.7 mm, and follows with a constant cross-section area test section 26H in length. The te st section is symmetric and has the option of optical acc ess through covering windows. It was water-cooled for combustion tests.

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40 Fuel was injected transverse to the airf low and into the reci rculation region 0.2H from the base of the step, or transverse to the airflow and at a distance 4H upstream of the step; a schematic diagram is shown in Figure 2-1 (b). For base injection, five 0.5-mm dia. holes equally spaced on each side of the test section wall were used. For upstream injection, two 1.0-mm dia. holes equally spa ced on each side of the isolator wall were used. Helium having molecular weight clos e to hydrogen, and argon having molecular weight close to propane, were injected as simulated fuel in non-reacting flow tests. Hydrogen was injected in combustion tests. A LabView program and associated National Instruments hardware was used to monitor experimental conditions such as ai rflow stagnation pressure, temperature and Mach number at the nozzle exit, fuel stagnati on pressure and temperature, test section wall static pressure and temperature dist ribution along the airflo w direction. Mass spectrometry (MS) and planar laser induced fl uorescence (PLIF) were used to determine the species mole fraction distribution in the recirculation region as described below. Mass Sampling and Analysis Hardware The physical location of mass sampling ports in the recircula tion region behind the step is shown in Figure 21 (b). The coordinate system is also shown in the figure. The test section window wall covering the st ep has five mass sampling ports in the recirculation region along the axial x-direction equally spaced from x/H = 0.5 to 3.5 and along y/H = 0.3. These ports are 0. 6 mm inner diameter steel tubes that end at the test section window wall and do not physically intrude into the re circulation region (z = 0). In separate tests, other tubes are inserted from the window wall to verify the twodimensionality of species distribution in the recirculation region. For non-reacting flow

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41 tests, three stainless steel tubes are placed at x/H = 2.0, y/H = 0.3, the location of port #3 in wall sampling; they penetrate into the test section to sample species at three different depths, equally spaced in the inflow z-direction from z/W = 0.33 to 1.0. Here W = 12.7 mm is the test section half-width. For com bustion tests, five 0.8 mm inner diameter ceramic tubes are inserted to z/W = 0.5 from the wall at the same axial locations as the wall sampling ports. Exposure to high temperatures of hydrogen combustion flow field in the flameholding region can cause metals to melt or oxidize. Hence the inflow sampling tubes for combustion tests were made of cer amics to withstand the high temperatures; inspection of these tubes after experiments showed that they remained intact during the tests. A schematic diagram of mass sampling from the recirculation region and subsequent, real-time analysis by a mass spect rometer is shown in Figure 2-2 (a). In combustion tests, the extracted species passe d through a water-cooled jacket on their way to the mass spectrometer to quench the reactions and freeze the species composition coming out of the combustion chamber. The j acket was supplied with cold water coming out of a chiller at 283 K. The temperatur e drop of the sampled mixture while passing through the cooling jacket resulted in condens ation of water vapor and much of it could not reach the mass spectrometer. Hence the corrected XH2O was deduced from the oxygen deficit in the product mixture. The sampling tubes coming out of the reci rculation region were connected to a manifold having six 0.6-mm diam eter input tubes and a sing le 1.8-mm diameter outlet tube connected to the mass spectrometer. Th e input of species to the manifold was regulated by a series of computer-controlled miniature solenoid va lves that supply gas

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42 mixture from one sampling port at a time for analysis. Sampling from each port was preceded by injection of nitrogen in the mani fold to purge the line and flush the species from the previous port, hence preventing mi xing of samples from two adjacent ports. A LabView program and associated National Inst ruments hardware was used to send analog signals to operate the miniature solenoid va lves in the desired sequence. The sampling time at each port and the purge time before each sampling were input to the software. The software is attached in Appendix A. Due to small volume of the sampling system, only a small mass of species was extracted from the recirculation region, which in turn facilitated quenching of the r eacting species and also helped the mass spectrometer attain steady state measurements almost instan taneously while switching from one port to another. The species were analyzed by Stanfo rd Research Systems RGA-300 mass spectrometer, shown in Figure 2-2 (b), that us es electron ionization to ionize the sampled gas, RF quadrupole filter to sort species according to their mass-to-charge ratio, and Faraday cup to detect ion currents. The inst rument is controlled and operated by software and associated electronics. The ionizer, filter and detector ar e enclosed in a clean vacuum chamber and require an opera ting pressure range of 10-4 torr (1.3 x 10-7 atm) to ultra high vacuum. Such low pressure is attained in two stages. In first stage, a rotary pump brings the inlet pressure down to about 60 mtorr (7.9 x 10-5 atm). In second stage, a diffusion pump and a rotary pump operate in series to bring the pres sure further down to vacuum conditions. The spectrometer can detect species up to a massto-charge ratio of 300 and has a resolution of 0.5 AMU @ 10% peak height The sensitivity factor of the instrument, defined as the signal detected per unit partial pressure of a given species (Amp / torr),

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43 varies for different gases. Hence calibrati on of the instrument was performed for the following gases: helium, nitrogen, oxygen and ar gon. The sensitivity factor of nitrogen is used as the baseline and sensitivity factor s of other gases are normalized with this baseline. The relative sensitivity factor (Sgas / SN2), also referred as the calibration factor, is shown in Table 2-1 for different gases. Data Processing The composition of gas in the recirculation region was analyzed by the mass spectrometer in partial pressu re of species vs. time mode. The species scanned were nitrogen (m/z=28,14), oxygen (m/z=32,16), helium (m/z=4), argon (m/z=40,20) for nonreacting experiments, and nitrogen (m/z=28,14), oxygen (m/z=32,16), hydrogen (m/z=2,1), water (m/z=18,17) for com bustion experiments. Sampling was done sequentially for 5 sec at the purge port and for 20 sec at each of the sampling ports. The spectrometer had a fast response time of about 2-3 sec to the change in composition of gas while switching from one port to another. The local mole fraction of a given species in the sample was determined from the part ial pressures of all the component species recorded by the mass spectrometer. The time-averaged fuel mole fraction at a port was obtained by averaging the mole fractions obtained over the sampling time period. The species mole fractions were corrected using calibration factors for individual gases. The background level of argon in the incoming ai rflow for non-reacting experiments, and the background levels of hydrogen and water for co mbustion experiments were subtracted to determine the actual mole fractions of these species. Figure 2-3 shows the species mole fraction distribution in the reci rculation region as the solenoid valves switch sequentially from one sampling port to another under st eady experimental condi tions; a non-reacting

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44 flow case with helium as the simulant fuel is shown in Figure 2-3 (a) and a reacting flow case with hydrogen combustion is shown in Figure 2-3 (b). The global mole fraction/equivalence ratio of fuel was determined from the total mole of fuel injected and the total mole of air traveling through th e test section. Both local and global mole fractions/equivalen ce ratios are indicated in the results. Optical Diagnostics Hardware Since mass spectrometry was limited to point-wise measurements in the flow, it was complemented by acetone planar laser in duced fluorescence (PLIF) to obtain fuel distribution in a 2D plane in the recirculation region for non-reacting flow. The plane of measurement was along z/W = 0.9. The PLIF measurements overlap with MS data at the point x/H = 2.0, y/H = 0.3, z/W = 0.9; Figure 2-1 (b) shows the laser sheet in the test section and the common point between MS a nd PLIF where measurements from the two techniques are compared in Chapter 3. A schematic diagram of PLIF setup is shown in Figure 2-4. The te st section used in PLIF measurements had step on only one side unlike on both sides for the test section used in mass spectrometer measurements. Howe ver, since the airflow arriving at the step base is supersonic, the flow field in the r ecirculation region is the same for both test sections. The test section side wall window was made of gla ss for laser sheet delivery in the recirculation region. The te st section front windows next to the step were also made of glass to provide visual access of the m easurement plane via a camera. Before being injected in the test section, the fuel was seeded with acetone va por by bubbling it into a chamber partially filled with liquid acetone at room temperature (295 K).

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45 A pulsed Nd:YAG Spectra-Physics laser wa s used for LIF excitation of acetone. The laser pulses at a frequency of 10 Hz and can have wavelength outputs of 1064, 532, 355 and 266 nm. The fourth harmonic 266 nm b eam was used for acetone PLIF. At 266 nm wavelength, the laser pulse had 70 mJ en ergy and a pulse width of 4-5 ns. The beam diameter coming out of the laser was 7 mm. It was converted into a sheet of light and delivered in the test section using a range of optics suited for UV range of laser light. The beam was first oriented towards the region of in terest in the test section using three 25.4 mm diameter mirrors. Then the beam diamet er was increased and collimated using spherical lenses. Two 25.4 mm diameter spheri cal lenses, a concave lens with f = -30 mm and a convex lens with f = 50 mm were placed next to each other with an effective focal length of f = -75 mm. The effective concave lens diverged the beam and increased its diameter. A 50.8 mm diameter spherical convex lens with f = 300 mm was placed 205 mm ahead of them. This resulted in a collimat ed beam with an increased diameter of 40 mm. This beam was then converted to a sheet of light using cylindrical lenses. Two 50.8 x 50.8 mm cylindrical lenses were used, a convex lens with f = 300 mm and a concave lens with f = -300 mm. The concave lens was placed 75 mm ahead of the convex lens, and together they formed an effective conve x lens, with the step placed 865 mm away at the focal point. The effective cylindrical le ns converged the beam in the horizontal direction while leaving it unchanged in th e vertical direction. The beam thus progressively converged into an ellipse with diminishing minor axis and unchanged major axis, and it formed a sheet 40 mm wide a nd less than 0.5 mm thick at the step. The sheet thickness did not change appreciably in the recirculation region due to high effective focal length of the lenses.

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46 For image acquisition of acetone fluoresce nce in the visible spectrum, a Cooke corporation intensified CCD camera and its associated software was used. The camera has a resolution of 1280 x 1024 pixels, 12-bit dyn amic range and shutter speed down to 3 ns. The software was used to set the camera parameters before taking images and also to save the captured images on the computer. Th e images were captured with an exposure time of 300 ms and a gain of 65%, w ith 100% representing the maximum gain achievable. A 2 x 2 binning was performed on the image in horizontal and vertical direction, that is, the average intensity of a square comprising 4 pixels was represented as the value for one pixel that replaced the 4 pixels. The camera resolution was 12 pixels/mm. Image Acquisition and Processing For each image acquisition, a total of 20 images were captured to get a timeaveraged image over 2 s. Three set of images were required for analyzing each experimental condition: the background image [Figure 2-5 (a)], the laser sheet profile image [Figure 2-5 (b)], and the actual experiment image. To obtain the laser sheet profile, the test section was closed and filled uniformly with acetone vapor, and then excited by laser sheet. Since the acetone concentration was uniform, the acquired image captured the spatial intensity variation in the laser sheet due to its Gaussian profile. The nature of the experiment created challenges in getting a low background image. The air arrived at the step base with M = 1.6 and had a low st atic temperature of T = 200 K. However, it slowed down in the recirculation region to low subsonic Mach numbers and the static temperature recovere d close to the stagnation temperature of 300 K. Hence a temperature gradient existed in the shear layer separating the recirculation region and the main airflow. This low air te mperature caused the acetone vapor injected

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47 in the test section to condense in the shear layer. The liquid acetone sprayed onto the glass window and created a high background fo r the camera. This issue was overcome by using a heat gun to heat the gl ass window up to 365 K just be fore starting airflow through the test section. Even as the heat gun wa s in operation during th e operation of the wind tunnel, the glass window temper ature dropped rapidly due to contact with cold airflow. Hence fuel seeded with acetone was injected in the test section and images were acquired within the first few minutes of starting the wind tunnel so that the glass window temperature would not drop below 329 K, th e boiling temperature for acetone. This procedure reduced the background quite s ubstantially. The background image set was acquired just after the experiment was over a nd the airflow and heat gun were turned off. The post processing of images was performed using a program written in Matlab. The software is attached in Appendix B. An image is handled by Matlab as a 2D matrix containing light intensity at each pixel, hence image processing is essentially an operation with matrices. Each set of laser sheet prof ile, background and actua l experiment images was averaged to get a time-averaged image. Then the background image was subtracted from the experiment image. The ultraviolet signal from Rayleigh scattering by acetone or from Mie scattering by condensed acetone drop lets was filtered by the camera lens since it transmits only the visible spectrum of light. The laser sheet profile image was used to correct the experiment image for spatial nonuniformity in laser intensity using the following formula. ) ( )] ( max[ ) ( ) ( p m I p m I n m I n m Ilaser laser d uncorrecte corrected (2-1) where I: pixel intensity m, n: arbitrary row and column location of a pixel in the image

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48 p: a fixed column in the laser sheet image With the exception of airflow shear layer a nd sonic fuel jet injected in the test section, the temperature distribut ion in the recirculation region is rather uniform. Further, the acetone LIF signal does not vary with temperature for low temperature range of 200300 K63, 64, 65. The pressure distribution in the reci rculation region is rather uniform. The acetone signal does not get quenched by oxygen62. Hence the LIF signal intensity variation in the corrected im age is independent of pressure, temperature and oxygen concentration; it varies only with the concen tration of acetone. With the assumption that acetone distribution in the flow is the same as that for the fuel, the fuel mole fraction distribution in the flow was determined quant itatively. The acetone LIF signal at the fuel injection location, where fuel mole fraction is 100 %, was taken as the reference point for other pixels in the image. The fuel mole fract ion at a given pixel was then determined by the ratio of LIF intensity at that pixel to the LIF intensity at the reference point. 100 (%)reference fuelI I X (2-2) For each experimental condition, image averaging was performed for 3 repeatability experiments. The fuel mole fraction distribution image, which does not depend on the laser intensity variation, was used to determine the average PLIF image and standard deviation for repeatability tests. The temporal variation in laser sheet profile was quantified for the 60 min duration over which repeatability experiments were performed. The image processing software is attached in Appendix B. Figure 2-6 shows the laser sheet profile at y/H = 1.1 for three time intervals after the laser was started. Also included is the profile when the laser was s hut down and restarted. The figure shows only a marginal change in laser profile over 85 min, the average standard deviation was 4%.

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49 Restarting the laser did not affect the profile. The spikes in the profile reflect the quality of the beam coming out of the laser. Convolution filtering was applied on the average PLIF image to smoothen the gradients due to noise. A 3 x 3 filter was app lied to the image, that is, the signal intensity at a given pixel was replaced by the average in tensity in a 3 x 3 pixel square with the given pixel at its center. The fuel mole frac tion distribution was obtained from the filtered image. The binary diffusion coeffici ent of a gas in air is i nversely proportio nal to the square root of its molecular weight; it indica tes that helium diffuses about 4 times faster in air compared to acetone and argon diffuses at about the same rate as acetone. Hence acetone will not be able to trace helium accurately while it will still trace argon reasonably well. It suggests that the fuel mole fraction measurements determined from acetone PLIF may be more accurate for argon as the injectant than helium. Condensation of acetone vapor into droplets may also affect the PLIF signal intensity.

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50 Table 2-1. Mass spectrometer calib ration factors for various gases Sgas / SN2 Helium Nitrogen Oxygen Argon 0.6 1.0 0.9 1.2

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51 Figure 2-1. Description of the test sec tion showing fuel injection and mass sampling locations: (a) image and (b) schematic diagram. Testsection No zzl e Isolator Upstream injection Base injection Mass sampling ports and cooling jacket (a) (b)

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52 Figure 2-2. Mass sampling from the recircula tion region behind the step for analysis by the mass spectrometer: (a) schematic diagram (b) image. (a) Solenoid valves Inlet Diffusion pump Turbo pump Vacuum chamber El ect r o ni cs (b)

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53 Figure 2-3. MS measurements of species mo le fraction distribution in the recirculation region as the solenoid valv es switch sequentially from one sampling port to another under steady experimental c onditions (a) non-reacting flow with helium as the simulant fuel (b) reac ting flow with hydrogen combustion. 0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 020406080100120140160time (sec)XHe (%) (a) 0 5 10 15 20 25 30 35 40 45 50 020406080100120140time (sec)X (%) H2 H2O O2 (b) H2 H2O O2

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54 Figure 2-4. Schematic diagram of planar laser-induced fluorescence (PLIF) setup.

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55 (a) (b) Figure 2-5. Sample PLIF image for (a) background (b) laser sheet.

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56 Figure 2-6. Temporal varia tion of laser sheet profile at y/H = 1.1.

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57 CHAPTER 3 RESULTS: NON-REACTING FLOW The study included non-reacting cases with the air conditions maintained at M = 1.6, T0air = 300 K, P0air = 4.8 atm. Helium and argon were injected as fuel simulants at two different pressures from the base of the step and from a location upstream of the step; the configuration is described in Chapter 2. The molecular weight of helium is close to hydrogen, while argon has molecular weight cl ose to propane. The data obtained from non-reacting flow tests is limited since com bustion of fuel-air mixture does not take place. However, it gives the fuel distributi on in the flameholding region just before ignition of fuel-air mixture. Based on this information from non-reac ting flow, the effect of various parameters, such as fuel inje ction pressure and location, on flameholding characteristics in actual combustion can be estimated. The mass exchange rate across the main airflow shear layer is higher for non -reacting flow compared to reacting flow24. For non-reacting flow, this will bring more air into the recirculation region, hence the flameholding region may be leaner in fuel compared to reacting flow. Mass Spectrometry (MS) Base Injection: Helium Non-reacting flow experiments were perfor med with helium as the fuel simulant. Each experiment was performed three times for repeatability. The average standard deviation in XHe was 4 %. Mass sampling of the re circulation region species was done along the wall in the x-direction with helium injected at the base of the step in P0air = 4.8 atm airflow. The wall pressure distribution for airflow without fuel injection is shown in

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58 Figure 3-1. The supersonic airflow expands at the step base, as indicated by the sharp drop in pressure at that loca tion. The shear layer formed due to separation of the airflow boundary layer at the step base is pushed towards the wall and it reattaches downstream of the step. An oblique shock is formed at the reattachment point, causing a pressure rise at that location. Based on the pressure rise in the plot, the shear layer reattachment point is estimated to be 1.5-2.0 H. Oblique shocks result in a pressure rise towards the end of test section to match the ambient atmospheri c pressure. The pressure distribution in the recirculation region remained unaffected due to fuel injection in the non-reacting flow tests. It indicates that fuel injection did not cause a substa ntial increase in the mass flow rate in the recirculation region. Helium was injected at two pressures, a moderate stagnation pressure P0He = 5.4 atm and a high stagnation pressure P0He = 12.0 atm. The wall distribution of XHe in the recirculation region for the two P0He is shown in Figure 3-2 (a ). The standard deviation bars are shown along with the average mole fr actions for the repeat ed experiments. The fuel injection location is indicated on the horiz ontal axis. It is observed that the fuel mole fraction decreases in the x-direction and away from the fuel injection location, especially for lower P0He. Increasing P0He substantially results in a corresponding increase in XHe in the recirculation region. The XHe distribution shows more non-uniformity at higher P0He. Mass sampling of the recirculation region species was done in the inflow zdirection for the same airflow and fuel in jection conditions as in wall sampling. The inflow sampling was done at x/H = 2.0, y/H = 0.3. The inflow distribution of XHe in the recirculation region is shown in Figure 3-2 (b) for the two P0He. At both pressures, the inflow XHe are much higher than the wall measured XHe, specifically, up to 4-5 times.

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59 However, XHe distribution away from the wall is ra ther uniform, indi cating a well mixed fuel-air mixture at th e axial location probed. Table 3-1 indicates the locally measured and the global XHe for wall and inflow sampling. The global XHe, defined in the earlier chapters, is obtained from the total moles of helium injected in the test section and th e total moles of air fl owing through the test section. For both P0He, the locally measured XHe is up to 3 times more than the global estimate for wall sampling and about 10 times more than the global estimate for inflow sampling. It shows that even as the global XHe suggests a fuel-lean mixture, a fuel rich mixture can exist in the recirculation region. Helium has a molecular weight close to hydrogen, hence the XHe distribution obtained from wall and inflow sa mpling is approximated to be similar to that of hydrogen injected under identical test conditions. Us ing this assumption, if hydrogen had been injected, the local equiva lence ratios of hydrogen ( H2) could be determined from the local mole fraction measurements. For P0H2 = 5.4 atm, the local H2 is estimated in the range of 0.04 0.4 for global H2 = 0.04. For P0H2 = 12.0 atm, the local H2 is estimated in the range of 0.1 0.7 for global H2 = 0.1. This confirms that a richer fuel composition mixture exists in the recirculation region even as the global H2 suggests a relatively fuellean mixture. It is expected since only a frac tion of the main airflow is entrained into the recirculation region. Hence the global H2 does not capture the actual non-premixed conditions existing in the flam eholding recirculation region. Base Injection: Argon Argon was injected at the base of the step for identical airflo w and fuel injection conditions as in the case of helium inject ion. The average standard deviation in XAr was 3 %. Mass sampling of the recirculation regi on species was done along the wall in the x-

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60 direction. The wall distribution of XAr in the recirculation region for the two P0Ar is shown in Figure 3-3 (a). The XAr distribution pattern is similar to that for helium injection, especially for low injection pressure. Mass sampling of the recirculation region species was done in the inflow zdirection for the same airflow and fuel in jection conditions as in wall sampling. The inflow distribution of XAr in the recirculati on region for the two P0Ar is shown in Figure 33 (b). As in wall sampling, the inflow XAr distribution has a pattern similar to that for helium injection [Figure 3-2 (b )], especially for low inj ection pressure. The inflow XAr measurements are up to 2-3 times higher than the wall measured XAr. The locally measured and the global XAr for wall and inflow sampling are shown in Table 3-2. The local to global Xfuel ratio can be interpreted as the fuel mole fraction in the recirculation region for a unit mole fraction of fuel injected in the test section. A comparison of data in Table 3-2 with Ta ble 3-1 shows that the local to global Xfuel ratio is higher for argon than helium, that is, highe r concentration of argon is found in the recirculation region than helium for a unit mole fraction of fuel injected in the test section. This could be attributed to diffusion; the binary diffusion coefficient of a gas in air is inversely proportional to the square root of its molecular weight, hence argon diffuses about 3 times slower in air compared to helium. Argon has a molecular weight close to propane, hence the XAr distribution obtained from wall and inflow sampling is approximated to be same as that of propane injected under identical test conditions. For P0C3H8 = 5.4 atm, the local C3H8 is estimated in the range of 0.5 2.5 for global C3H8 = 0.1. For P0C3H8 = 12.0 atm, the local C3H8 is estimated in the range of 1.0 4.3 for global C3H8 = 0.2. This shows that a fuel-rich

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61 mixture exists in the recircul ation region even as the global C3H8 suggests a fuel-lean mixture. Upstream Injection: Helium Helium was injected upstream of the step in identical airflow c onditions as the base injection experiments. The av erage standard deviation in XHe was 6 %. Helium was injected at two pressures, P0He = 2.4 atm and 5.1 atm. The corresponding dynamic pressure ratios [Pdynamic = HeVHe 2 / airVair 2] are 0.5 and 1.0 respectively. Mass sampling of the recirculation region speci es was done along the wall in x-direction and in the inflow z-direction. The wa ll distribution of XHe in the recirculation region for the two P0He is shown in Figure 3-4 (a) and the inflow di stribution is shown in Figure 3-4 (b). The inflow sampling shows an al most proportional increase in XHe with P0He. Table 3-3 summarizes the local and global XHe for wall and inflow sampling. The plots and the table show quite low levels of XHe in the recirculation region, hence indicating that upstream injection of helium is not effective in supplying fuel to the recircul ation region. In fact, the local XHe is lesser than global XHe, indicating that the light gas penetrates through the main airflow shear layer and only a small quantity reaches the recirculation region. The XHe distribution obtained in the experiment s is approximated to be the same as that of hydrogen injected under identical test conditions. For P0H2 = 2.4 atm, the local H2 is estimated in the range of 0.01 0.02 for global H2 = 0.02. For P0H2 = 5.1 atm, the local H2 is estimated in the range of 0.01 0.03 for global H2 = 0.05. The fuel-lean conditions for upstream injection mode s uggest difficulty in flameholding in a combustion experiment with hydrogen as the fuel.

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62 Upstream Injection: Argon Argon was injected upstream of the step fo r identical airflow and fuel injection conditions as in case of helium injection. Hence, argon was injected at the same dynamic pressure ratios as helium. The average standard deviation in XAr was 3 %. Mass sampling of the recirculation region speci es was done along the wall in x-direction and in the inflow z-direction. The wa ll distribution of XAr in the recirculation region for the two P0Ar is shown in Figure 3-5 (a) and the inflow distribution is show n in Figure 3-5 (b). Table 34 summarizes the local and global XAr for wall and inflow sampling. Unlike the case of helium, the heavier gas argon reach es the recirculation region in quantities larger than the global XAr. It can be seen from the plot s and from the local to global XAr ratios in the table that an increase in the upstream P0Ar does not result in a corresponding increase in the amount of fuel reaching the recirculation regi on. If fuel is inject ed upstream at a low dynamic pressure ratio, it seep s into the boundary layer of the incoming airflow which carries it into the recirculation region. If the upstream fuel injection dynamic pressure ratio is increased, part of the fuel pene trates through the airf low boundary layer and escapes into the core airflow; hence le ss fuel reaches the recirculation region. The XAr distribution obtained in the experiment s is approximated to be the same as that of propane injected under identical test conditions. For P0C3H8 = 2.4 atm, the local C3H8 is estimated in the range of 0.2 0.3 for global C3H8 = 0.06. For P0C3H8 = 5.1 atm, the local C3H8 is estimated in the rang e of 0.2 0.3 for global C3H8 = 0.1. The fuel-lean conditions in the recirculat ion region for upstream injection mode, along with much lower flameholding limits for hydrocarbons as compared to hydroge n suggest difficulty in flameholding in a combustion expe riment with propane as the fuel.

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63 Planar Laser Induced Fluorescence (PLIF) Base Injection: Helium PLIF imaging of the recirculati on region fuel distribution in the z/W = 0.9 plane was performed for identical airflow and fuel injection conditions as the mass sampling experiments. Each experiment was perfor med 3 times for repeatability. The average standard deviation in XHe, after deducting the XHe deviation due to th e temporal variation in laser sheet profile discussed in Chapter 2, was 8 %. The results for helium injection are shown in Figures 3-6 and 3-7 for P0He = 5.4 atm and 12.0 atm respectively. The PLIF image for P0He = 5.4 atm is shown in Figure 3-6 (a) and the XHe distribution is shown in Figure 3-6 (b). The expansion of airflow at the step pushes the shear layer towards the test section wall. The fuel injection holes are inclined relative to the step base. This is clearly visible in the figures as the fuel je t impinges on the step base and forms a plume above it. This fuel injection configuration helps the fuel remain and mix within the recirculation region. The PLIF image for P0He = 12.0 atm is shown in Figure 3-7 (a) and the XHe distribution is shown in Figure 3-7 (b). In agreement with the mass spectrometer measurements, the fuel remains in the recirculation region even as the injection pressure is increased. The XHe distribution shows more non un iformity for higher injection pressure. The global XHe for P0He = 5.4 atm and 12.0 atm are 1.1 % and 2.5 % respectively. As in mass spectrometer measurements, the local XHe in the recirculation region and shear layer [Figure 3-6 (b) and 3-7 (b)] are an order of magnit ude higher than the corresponding global XHe values. The XHe distribution obtained in the experiments is approximated to be the same as that of hydroge n injected under identi cal test conditions. For P0H2 = 5.4 atm, the global H2 = 0.04 and for P0H2 = 12.0 atm, the global H2 = 0.1.

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64 The global H2 values suggest a fuel lean mixture fo r both injection pressures. With the observation that H2 = 1 corresponds to 30 % XH2 in a hydrogen-air mixt ure, it is seen in Figure 3-6 (b) that for P0H2 = 5.4 atm, part of the recirc ulation region has a fuel-rich mixture with H2 > 1. Figure 3-7 (b) shows that for P0H2 = 12.0 atm, the entire recirculation region, excluding the main airflow shear layer, has H2 > 1. Base Injection: Argon Argon was injected at the base of the step for identical airflo w and fuel injection conditions as in the case of helium inject ion. The average standard deviation in XAr, after deducting the XAr deviation due to the temporal varia tion in laser sheet profile, was 8 %. The results for ar gon injection are shown in Figures 3-8 and 3-9 for P0Ar = 5.4 atm and 12.0 atm respectively. The PLIF image for P0Ar = 5.4 atm is shown in Figure 3-8 (a) and the XAr distribution is shown in Figure 3-8 (b). The PLIF image for P0Ar = 12.0 atm is shown in Figure 3-9 (a) and the XAr distribution is shown in Figure 3-9 (b). As in the case of helium injection, the fuel remains within the recirculation region for both injection pressures. By comparing fuel mole fraction distribution in Figure 38 (b) with Figure 3-6 (b) for low injection pressure and Figure 3-9 (b) with Figure 3-7 (b) for high injection pressure, it is observed that argon distribution in the recirc ulation region is similar to helium, especially for low injection pressu re. This is in agreement with the mass spectrometer measurements. The global XAr for P0Ar = 5.4 atm and 12.0 atm are 0.4 % and 0.8 % respectively. As in helium injection, a much richer fuel composition exists in the recirculation region [Figure 3-8 (b) and 3-9 (b)] than suggested by global XAr values. The XAr distribution obtained in the experiments is approximated to be the same as that of propane injected under identical test conditions. For P0C3H8 = 5.4 atm, the global C3H8 = 0.1 and for P0C3H8

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65 = 12.0 atm, the global C3H8 = 0.2. The global C3H8 values suggest a fuel lean mixture for both injection pressures. With the observation that C3H8 = 1 corresponds to 4 % XC3H8 in a propane-air mixture, it is seen in Figures 3-8 (b) and 3-9 (b) that for both injection pressures, the entire recirculation region and the main airflow shear layer is fuel-rich in composition, with C3H8 > 2. Upstream Injection PLIF imaging of the recirculation regi on fuel distribution was performed for identical airflow and fuel in jection conditions as the mass sampling experiments. For all test conditions, the PLIF signal was barely noticeable above the background. It is in agreement with mass spectrometer measurements and shows that hardly any fuel reaches the recirculation region fo r upstream injection. Comparison between MS and PLIF data The fuel mole fraction measurements at x/H = 2.0, y/H = 0.3, z/W = 0.9 obtained from PLIF and MS are compared in Figure 310 for base fuel injection of helium [Figure 3-10 (a)] and argon [Figure 3-10 (b)]. For PLIF the fuel mole fractions obtained from unfiltered image and averaged over the mass sampling tube cross-section are reported. Filtering increased the fuel mole fraction by 10 % for helium and 2 % for argon injection, hence unfiltered data was used for comparison. The PLIF and MS data do not overlap with each other for all test conditions. One of the reasons for the difference between PLIF a nd MS measurements is the intrusive nature of mass sampling for MS measurements. The outer diameter of th e 3 inflow sampling tubes was 0.9 mm; the small diam eter tubes were selected to minimize disturbances to the flow field. However, changes in the local flow fiel d due to the presence of sampling tubes cannot be eliminated completely. A more signi ficant disturbance in the local flow field

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66 was caused by suction applied at the sampling port tip to extract gas mixture from the recirculation region into the mass spectromete r; the pressure grad ient was high, with about 0.5 atm in the flow field behind the st ep and near vacuum at the entrance of the mass spectrometer. Hence, a difference between MS and PLIF data is expected due to different local flow fields in the two measurements. As discussed in Chapter 2, the limitation of PLIF measurements is that acetone is expected to trace argon better than helium due to the difference in diffusion rates. Acetone diffuses about 4 times slower than helium and about the same rate as argon. Hence acetone PLIF measurements will over-es timate helium mole fraction in the flow. This is observed in the MS-PLIF data comp arison; the difference between acetone PLIF fuel mole fraction measurements and corresponding MS measurements is higher for helium compared to argon. The difference could also be due to a slig ht misalignment of the measurement point between the two techniques due to shaking of the test section as air flowed through it. The point at which PLIF and MS data are compar ed lies in the airflow shear layer. In this region, sharp gradients exist in the x-y plane, even though the flow is rather uniform in the z direction as seen in inflow MS measur ements. The lateral test section movement observed by the camera during PLIF tests was 3 pixels ( y / H = 0.02). For P0He = 5.4 atm and 12.0 atm test conditions, a lateral shift in measurement location from the reference point at y/H = 0.3 by y / H = 0.02 results in a 3 % change in fuel mole fraction. Hence shaking of the test section is not a significant reason for the difference between MS and PLIF data.

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67 Table 3-1. Base fuel in jection: global and local He. Table 3-2. Base fuel injection: global and local Ar. Table 3-3. Upstream fuel injection: global and local He. Table 3-4. Upstream fuel injection: global and local Ar. P0 He (atm) local He (%) global He (%) local / global He Wall sampling 5.4 1.1-3.8 1.1 0.9-3.4 12.0 2.6-8.3 2.5 1.0-3.4 Inflow sampling 5.4 2.2-11.8 1.1 1.9-10.4 12.0 5.7-21.5 2.5 2.3-8.7 P0 Ar (atm) local Ar (%) global Ar (%) local / global Ar Wall sampling 5.4 2.0-5.0 0.4 5.5-13.9 12.0 3.8-9.9 0.8 4.9-12.7 Inflow sampling 5.4 2.9-9.9 0.4 7.9-27.4 12.0 5.2-17.1 0.8 6.6-21.9 P0 He (atm) local He (%) global He (%) local / global He Wall sampling 2.4 0.2-0.3 0.7 0.3-0.5 5.1 0.3-0.4 1.5 0.2-0.3 Inflow sampling 2.4 0.3-0.6 0.7 0.4-0.9 5.1 0.4-1.0 1.5 0.2-0.7 P0 Ar (atm) local Ar (%) global Ar (%) local / global Ar Wall sampling 2.4 0.7-0.9 0.2 3.3-4.0 5.1 0.8-1.0 0.5 1.7-2.0 Inflow sampling 2.4 0.7-1.0 0.2 3.3-4.5 5.1 0.9-1.3 0.5 1.9-2.8

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68 Figure 3-1. Wall pressure distribution for non-reacting flow. P0air = 4.8 atm, Mair = 1.6. The axial origin is placed at the step. 0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 -20-15-10-50510152025x/HPwall / Pnozzle exit

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69 Figure 3-2. Base fuel injection: MS measur ement of helium mole fraction distribution in the recirculation region for (a) wa ll sampling (b) inflow sampling. P0air = 4.8 atm, Mair = 1.6. 0 5 10 15 20 250.00.51.01.52.02.53.03.54.0x/HXHe (%) P0He=5.4atm P0He=12.0atm (a) 0 5 10 15 20 25 0.00.20.40.60.81.0z/WXHe (%) P0He = 5.4atm P0He = 12.0atm (b) P0He = 5.4 atm P0He = 12.0 atm P0He = 5.4 atm P0He = 12.0 atm

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70 Figure 3-3. Base fuel inject ion: MS measurement of argon mole fraction distribution in the recirculation region for (a) wa ll sampling (b) inflow sampling. P0air = 4.8 atm, Mair = 1.6. 0 5 10 15 200.00.51.01.52.02.53.03.54.0x/HXAr (%) P0Ar = 5.4atm P0Ar = 12.0atm (a) 0 5 10 15 20 0.00.20.40.60.81.0z/WXAr (%) P0Ar = 5.4atm P0Ar = 12.0atm (b) P0Ar = 5.4 atm P0Ar = 12.0 atm P0Ar = 5.4 atm P0Ar = 12.0 atm

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71 Figure 3-4. Upstream fuel injection: MS measurement of helium mole fraction distribution in the recirculation re gion for (a) wall sampling (b) inflow sampling. P0air = 4.8 atm, Mair = 1.6. 0.0 0.2 0.4 0.6 0.8 1.0 1.20.00.51.01.52.02.53.03.54.0x/HXHe (%) P0He = 2.4atm P0He = 5.1atm (a) 0.0 0.2 0.4 0.6 0.8 1.0 1.2 0.00.20.40.60.81.0z/WXHe (%) P0He = 2.4atm P0He = 5.1atm ( b ) P0He = 2.4 atm P0He = 5.1 atm P0He = 2.4 atm P0He = 5.1 atm

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72 Figure 3-5. Upstream fuel injection: MS measurement of argon mole fraction distribution in the recirculation re gion for (a) wall sampling (b) inflow sampling. P0air = 4.8 atm, Mair = 1.6. 0.0 0.5 1.0 1.5 2.00.00.51.01.52.02.53.03.54.0x/HXAr (%) P0Ar = 2.4atm P0Ar = 5.1atm (a) 0.0 0.5 1.0 1.5 2.00.00.20.40.60.81.0z/WXAr (%) P0Ar = 2.4atm P0Ar = 5.1atm (b) P0Ar = 2.4 atm P0Ar = 5.1 atm P0Ar = 2.4 atm P0Ar = 5.1 atm

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73 Figure 3-6. PLIF measurement for ba se injection of helium (a) image (b) XHe distribution (%). P0He = 5.4 atm, P0air = 4.8 atm, Mair = 1.6. (a) (b) Airflow Fuel 0.2H

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74 (a) (b) Figure 3-7. PLIF measurement for ba se injection of helium (a) image (b) XHe distribution (%). P0He = 12.0 atm, P0air = 4.8 atm, Mair = 1.6. Airflow 0.2H Fuel

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75 Figure 3-8. PLIF measurement for ba se injection of argon (a) image (b) XAr distribution (%). P0Ar = 5.4 atm, P0air = 4.8 atm, Mair = 1.6. (a) (b) Fuel Airflow 0.2H

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76 Figure 3-9. PLIF measurement for ba se injection of argon (a) image (b) XAr distribution (%). P0Ar = 12.0 atm, P0air = 4.8 atm, Mair = 1.6. (a) (b) Fuel Airflow 0.2H

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77 Figure 3-10. Comparison between MS and PLIF data for base fuel injection of (a) helium (b) argon. 0 5 10 15 20 25 30 0.00.20.40.60.81.0z/WXHe (%) P0He = 5.4atm (MS) P0He = 12.0atm (MS) P0He = 5.4 atm (PLIF) P0He = 12.0 atm (PLIF) (a) 0 5 10 15 20 0.00.20.40.60.81.0z/WXAr (%) P0Ar = 5.4 atm (MS) P0Ar = 12.0 atm (MS) P0Ar = 5.4 atm (PLIF) P0Ar = 12.0 atm (PLIF) (b) P0He = 5.4 atm (MS) P0He = 12.0 atm (MS) P0He = 5.4 atm (PLIF) P0He = 12.0 atm (PLIF) P0Ar = 5.4 atm (MS) P0Ar = 12.0 atm (MS) P0Ar = 5.4 atm (PLIF) P0Ar = 12.0 atm (PLIF)

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78 CHAPTER 4 RESULTS: REACTING FLOW The study included reacting cases with the air conditions maintained at M = 1.6, T0air = 300 K, P0air = 4.5 atm. Hydrogen was injected as fuel at two different pressures from the base of the step and from a locati on upstream of the step ; the configuration is described in Chapter 2. The re sults are described below. Base Injection: Hydrogen Combustion experiments were performed with hydrogen injected at the base of the step. Mass sampling of the recirculation re gion species was done along the wall in the xdirection. Hydrogen was injected at two stagnation pressures: P0H2 = 4.5 atm and P0H2 = 8.2 atm, the air stagnation pressure was P0air = 4.5 atm. These corresponded to global H2 of 0.04 and 0.08 respectively. The wall pressure distribution for the two P0H2 is shown in Figure 4-1. For P0H2 = 4.5 atm, the expansion of airflo w approaching the step is reduced due to heat released from combustion and pr essure rise at the step base. Hence the recirculation region length is more than the corres ponding non-reacting case. For P0H2 = 8.2 atm, the rather uniform pressure distributi on due to more heat release implies an even longer recirculation region and probably no reat tachment point for the shear layer. Figure 4-2 (a) shows the wall distribution of local H2 in the recirculation region for P0H2 = 4.5 atm. Figure 4-3 (a) shows the corresponding plot for P0H2 = 8.2 atm. The local H2 was deduced from the mole fractions of hydrogen and water in the product mixture. The fuel injection location is indicated on the horizontal axis in th e plots. A highly non-uniform H2 distribution is observed in the recircul ation region with a maximum around 2.2H. The

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79 local H2 goes up to 0.7 at P0H2 = 4.5 atm and up to 1.3 at P0H2 = 8.2 atm, showing a proportional increase due to an increase in P0H2. The wall distribution of products from hydrogen combustion for the two P0H2 is shown in Figures 4-2 (b) and 4-3 (b). XN2 has not been included in the plots. Concurrent with the earlier observation of a fuel-rich mixt ure existing in the reci rculation region, the combustion product composition shows a sign ificant proportion of unburned hydrogen. XH2O increases as we move downstream of th e step. However a significant proportion of unburned oxygen and only a small proportion of wa ter are a reflection of the limitation of sampling at the wall where the combusti on radicals get quenched and hence the composition can be quite different from elsewhere in the flow. Inflow mass sampling of the recirculat ion region species was done for the same airflow and fuel injection condi tions as in wall sampling. Figures 4-4 and 4-5 show the distribution of local H2 and product mole fractions in the recirculation region for P0H2 = 4.5 atm and P0H2 = 8.2 atm respectively. The experimental conditions were not identical over the repeated tests, both P0air and P0H2 varied by 0.3 atm. Hence the high standard deviations observed in the plots could be attr ibuted to the change in recirculation region mixture composition due to changes in airflo w and fuel injection conditions. The local H2 distributions in Figures 4-4 (a) and 4-5 (a ) show a decreasing amount of fuel as we go downstream of the fuel injection location. The local H2 does not increase proportionally with the increase in P0H2. As P0H2 increased, the local H2 increased unevenly in the recirculation region; the region close to the injection location experienced a lower increase in H2 than the region farther away in the x-direction. Thus, increasing P0H2 led to a reduction of local gradients in the recirculation region. The product mole fraction

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80 distributions in Figures 4-4 (b ) and 4-5 (b) show a fuel rich mixture in the recirculation region with plenty of unbur ned hydrogen and almost no oxygen. The proportion of unburned hydrogen drops rapidly as we go downs tream in the recirculation region. For low P0H2, once the hydrogen was completely consumed, the oxygen mole fraction increased. For high P0H2, hydrogen was still present at th e last sampling port and oxygen was virtually nonexistent. The temperature dr op of the sampled mixture while passing through the cooling jacket resulted in condens ation of water vapor and much of it could not reach the mass spectrometer. Hence the corrected XH2O was deduced from the oxygen deficit in the product mixture. Both corrected and uncorrected XH2O are plotted in the figures. Unlike the wall sampling experiment s, significant amount of water was produced and XH2O increases as we go downstream in the recirculation region. The local and global H2 for the two P0H2 obtained from wall and inflow samplings are compared in Table 4-1. As in the non-reacting flow test results discussed earlier, more fuel is observed away from the wall. For both P0H2, the local H2 is an order of magnitude higher than the suggested global value. Upstream Injection: Hydrogen Hydrogen was injected upstream of the step for identical airflow conditions as base injection. Hydrogen was injected at two stagnation pressures: P0H2 = 2.5 atm and P0H2 = 8.2 atm. The corresponding dynamic pressure ratios [Pdynamic = H2VH2 2 / airVair 2] are 0.5 and 1.6 respectively. However, a flam e could not be established for both P0H2, which is in line with the predictions of non-reacting flow experiments with helium injected upstream of the step. Mass sampling of the recirc ulation region species along the wall in xdirection showed a fuel-lean mixture with no water formed as a byproduct of combustion.

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81 A change in the injection configuration, e.g ., number of orifices, angled injection, etc., may lead to possibly holding the flame.

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82 Table 4-1. Base fuel in jection: global and local H2 P0 H2 (atm) local H2 global H2 local/global H2 Wall sampling 4.5 0.1-0.7 0.04 1.8-18.3 8.2 0.2-1.3 0.08 2.6-16.4 Inflow sampling 4.5 0.8-2.7 0.04 21.0-66.5 8.2 1.5-2.8 0.08 18.5-35.0

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83 Figure 4-1. Wall pressure distribu tion for hydrogen combustion tests. P0air = 4.5 atm, Mair = 1.6. 0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4-20-15-10-50510152025x/HPwall / Pnozzle exit P0H2=4.5atm P0H2=8.2atm P0H2 = 4.5 atm P0H2 = 8.2 atm

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84 Figure 4-2. Base fuel injection: Wall sampling results for (a ) hydrogen equivalence ratio distribution in the recirculatio n region (b) combustion species mole fraction distribution. P0H2 = 4.5 atm, P0air = 4.5 atm, Mair = 1.6. 0.0 0.5 1.0 1.5 2.0 2.5 3.0 0.00.51.01.52.02.53.03.54.0x/HH2 (a) 0 10 20 30 40 50 60 70 80 0.00.51.01.52.02.53.03.54.0x/HX (%) H2 H2O O2 (b) H2 H2O O2

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85 Figure 4-3. Base fuel injection: Wall sampling results for (a ) hydrogen equivalence ratio distribution in the recirculatio n region (b) combustion species mole fraction distribution. P0H2 = 8.2 atm, P0air = 4.5 atm, Mair = 1.6. 0.0 0.5 1.0 1.5 2.0 2.5 3.0 0.00.51.01.52.02.53.03.54.0x/HH2 (a) 0 10 20 30 40 50 60 70 80 0.00.51.01.52.02.53.03.54.0x/HX (%) H2 H2O O2 (b) H2 H2O O2

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86 Figure 4-4. Base fuel injection: Inflow sampling results for (a) hydrogen equivalence ratio distribution in the recirculatio n region (b) combustion species mole fraction distribution. P0H2 = 4.5 atm, P0air = 4.5 atm, Mair = 1.6. 0.0 0.5 1.0 1.5 2.0 2.5 3.0 0.00.51.01.52.02.53.03.54.0x/HH2 (a) 0 10 20 30 40 50 60 70 80 0.00.51.01.52.02.53.03.54.0x/HX (%) H2 H2O O2 H2O uncorrected (b) H2 H2O O2 H2O uncorrected

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87 Figure 4-5. Base fuel injection: Inflow sampling results for (a) hydrogen equivalence ratio distribution in the recirculatio n region (b) combustion species mole fraction distribution. P0H2 = 8.2 atm, P0air = 4.5 atm, Mair = 1.6. 0.0 0.5 1.0 1.5 2.0 2.5 3.0 0.00.51.01.52.02.53.03.54.0x/HH2 (a) 0 10 20 30 40 50 60 70 80 0.00.51.01.52.02.53.03.54.0x/HX (%) H2 H2O O2 H2O uncorrected (b) H2 H2O O2 H2O uncorrected

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88 CHAPTER 5 CONCLUSIONS Mass spectrometry (MS) and planar laser i nduced fluorescence (PLIF) were used to determine the species concentration distributi on in the flameholding recirculation region and free shear layer formed behind a recta ngular step in supers onic flow. Non-reacting and combustion tests were conducted and fuel related parameters such as the injection location, injection pressure and fuel type were varied. The conclusions are summarized below. Fuel injection location: Fuel injection at the base of the step wa s effective in supplying fuel directly into the flameholding region. Stable flames were achieved in combustion tests. Fuel injection upstream of the step was not effective in supplying sufficient amount of fuel to the recirc ulation region, hence a flame could not be sustained in combustion tests. An injection configur ation change, such as increasing the number of injection holes, could improve flameholding for upstream injection. Fuel concentration in th e recirculation region: Base injection: The local fuel concentration in the recirculation region was an order of magnitude higher than the suggested globa l fuel mole fraction since only a small part of the main airflow entered the re circulation region. For combustion tests with hydrogen, the recirculation regi on was predominantly fuel-rich in composition even for quite low global H2.

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89 The above observation implies that for the same global equivalence ratio, the flameholding region is richer in fu el composition for non-premixed case compared to premixed case. As the global equivalence ratio is increased, the nonpremixed case will flood the recirculation region with fuel and hence have worse fuel-rich flame stability limit than the premixed case. On the other hand, as the global equivalence ratio is reduced, the non-premixed case will still have sufficient fuel in the recirculation re gion and hence have a better fuel-lean stability limit than the premixed case. Non-reacting flow expe riments indicated a leaner composition of the recirculation region than the combustion tests. It could be due to higher air entrainment rate into the recirculation re gion through the main airflow shear layer for non-reacting flow tests than combustion tests. For non-reacting and reacting flow te sts, higher fuel concentration was measured inflow in the recirculation region than at the test section wall. Fuel injection pressure: Base injection: For non-reacting flow tests, fuel remain ed in the recirculation region and shear layer for both injection pressures. The fuel injection holes are inclined relative to the step base; such a configuration helps the fuel remain and mix within the recirculation region. Fuel distribution in the recirculation region was more nonuniform for higher fuel injection pressure. For combustion tests, increasing the fuel injection pressure resulted in reducing the fuel distribution gr adient inflow in the recirculation region.

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90 Upstream injection: The lighter gas, i.e. helium, penetrat ed through the airflow boundary layer along the wall for both injection pressures; the fu el quantity detected in the recirculation region was less than the suggested global value. At higher injection pressure for the heav ier gas, i.e. argon, the jet penetrated through the airflow boundary layer and less of it was carried into the recirculation region. However, for both argon injection pr essures, more fuel was detected in the recirculation region than th e indicated global value. Fuel type: Base injection: The fuel distribution pattern in the r ecirculation region was similar for helium and argon. However, argon concentration in the recirculation region was higher than helium for a unit mole fraction of fuel in jected in the test section. It could be due to slower diffusion rate of argon in air than that of helium in air. The above observation implies that as th e fuel injection pre ssure is increased, the heavier fuel will have higher fuel mole fraction in the recirculation region, and hence have a worse fuel-rich flame stability limit, than the lighter fuel. For the same reason, as the injection pressure is decreased, the heavier fuel will still have enough fuel mole fraction in the recirculati on region, and hence have a better fuellean flame stability limit, compared to the lighter fuel. This analysis is in addition to the fact that a heavier fuel such as propane has a much smaller flame stability curve than a lighter fuel such as hydrogen.

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91 Comparison of MS and PLIF data: The data from MS and PLIF at the co mmon point in the recirculation region do not overlap with each other. This is in pa rt due to the intrus ive nature of mass sampling for MS measurements; the lo cal flow fields in MS and PLIF experiments are not identical. Further, acetone PLIF over-estimates fuel mole fraction for helium compared to argon.

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92 APPENDIX A LABVIEW PROGRAM FOR MASS SAMPLING SEQUENCING Front Panel 15000 No of samples 5.0 Signal 0.0 Purge phase 0.0667 0 Purge frequency Hz 33.33 Purge duty cycle (%) 200.0 Sampling rate Hz STOP 14.29 Sampling duty cycle (%) 5.0 0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 75.0 0.0 10.0 20.0 30.0 40.0 50.0 60.0 70.0 Purge Sample 1 Sample 2 Sample 3 Sample 4 Sample 5 0.0143 0 Sampling frequency Hz 5.0 Purge time (s) 10.0 Sampling time (s) Purge Sample 1 Sample 2 Sample 3 Sample 4 Sample 5

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93 Block Diagram 0.00 0.00 True Pur g e dut y c y cle ( % ) No of sam p les Si g nal am p litude Pur g e fre q uenc y Sam p lin g rate Pur g e 0.0 s Pur g e p hase 0 [ 0..3 ]

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94 0.00 False Pur g e dut y c y cle ( % ) No of sam p les Si g nal am p litude Pur g e fre q uenc y Sam p lin g rate Pur g e 0.0 s Pur g e p hase 0 [ 0..3 ]

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95 -360 -360 4 Sam p lin g time ( s ) Pur g e time ( s ) 1 [ 0..3 ]

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96 0.00 0.00 True Sam p lin g dut y c y cle ( % ) Sam p lin g fre q uenc y Sam p lin g -1 0.0 s fre q uenc y chart-s 0.00 0.00 True Sam p lin g -2 0.0 s 0.00 0.00 True Sam p lin g -3 0.0 s 0.00 0.00 True Sam p lin g -4 0.0 s 0.00 0.00 True Sam p lin g -5 0.0 s 2 [ 0..3 ]

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97 0.00 False Sam p lin g dut y c y cle ( % ) Sam p lin g fre q uenc y Sam p lin g -1 0.0 s fre q uenc y chart-s 0.00 False Sam p lin g -2 0.0 s 0.00 False Sam p lin g -3 0.0 s 0.00 False Sam p lin g -4 0.0 s 0.00 False Sam p lin g -5 0.0 s 2 [ 0..3 ]

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98 1 ob0!sc1!md3!ch0 Pur g e 0.00 1 ob0!sc1!md3!ch1 0.00 Sam p le 1 1 ob0!sc1!md3!ch2 0.00 Sam p le 2 1 ob0!sc1!md3!ch3 Sam p le 3 0.00 1 ob0!sc1!md3!ch4 Sam p le 4 0.00 1 ob0!sc1!md3!ch5 Sam p le 5 0.00 3 [ 0..3 ]

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99 APPENDIX B MATLAB PROGRAM FOR PLIF IMAGE PROCESSING imbatchread.m function [immean]=imbatchread(flnhdr) %time averaging of a set of images d=dir([flnhdr '*.tif']); nFiles=length(d); disp(['reading flnhdr batch num2str(nFiles) files']); imtot=double(imread(d(1).name)); for i=2:nFiles, im=double(imread(d(i).name)); imtot=imtot+im; if floor(i/10)==i/10, disp(['read num2str(i) files']); end; end; immean=imtot/nFiles; %image axes origin shift from top-left to bottom-left corner [m,n]=size(immean); for i=1:m/2, temp=immean(m-i+1,:); immean(m-i+1,:)=immean(i,:);

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100 immean(i,:)=temp; end; imageprocessing.m %imageprocessing.m %Image processing for DiCamPro image batches laser='110805a'; %file name header of laser sheet file to use run_A='110505h'; %file name header of experiment file to use bgnd_A='110505i'; %file name header of background file to use run_B='110505j'; %file name header of experiment file to use bgnd_B='110505k'; %file name header of background file to use run_C='110505l'; %file name header of experiment file to use bgnd_C='110505m'; %file name header of background file to use %time averaging of a set of images [lasermean]=imbatchread(laser); [runmean_A]=imbatchread(run_A); [bgndmean_A]=imbatchread(bgnd_A); [runmean_B]=imbatchread(run_B); [bgndmean_B]=imbatchread(bgnd_B); [runmean_C]=imbatchread(run_C); [bgndmean_C]=imbatchread(bgnd_C); %background subtraction runcorr1_A=runmean_A-bgndmean_A; runcorr1_B=runmean_B-bgndmean_B;

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101 runcorr1_C=runmean_C-bgndmean_C; %laser sheet profile correction [m,n]=size(runcorr1_A); for i=1:n, runcorr2_A(:,i)=runcorr 1_A(:,i)./lasermean(:,170)*max(lasermean(:,170)); end; [m,n]=size(runcorr1_B); for i=1:n, runcorr2_B(:,i)=runcorr 1_B(:,i)./lasermean(:,170)*max(lasermean(:,170)); end; [m,n]=size(runcorr1_C); for i=1:n, runcorr2_C(:,i)=runcorr 1_C(:,i)./lasermean(:,170)*max(lasermean(:,170)); end; %fuel mole fraction distribution %I1=861;%80psi %I2=965;%80psi %I3=905;%80psi I1=669;%176psi I2=718;%176psi I3=523;%176psi relconc_A=100*runcorr2_A/I1; relconc_B=100*runcorr2_B/I2;

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102 relconc_C=100*runcorr2_C/I3; %plotting figure(11); Z=relconc_A; contour(Z,10); [C,h]=contour(Z,10); clabel(C,h,'manual'); axis xy; axis equal; axis tight; figure(12); image(runcorr2_A/10); colormap gray; axis xy; axis equal; axis tight; axis off; figure(21); Z=relconc_B; contour(Z,10); [C,h]=contour(Z,10); clabel(C,h,'manual'); axis xy;

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103 axis equal; axis tight; figure(22); image(runcorr2_B/10); colormap gray; axis xy; axis equal; axis tight; axis off; figure(31); Z=relconc_C; contour(Z,10); [C,h]=contour(Z,10); clabel(C,h,'manual'); axis xy; axis equal; axis tight; figure(32); image(runcorr2_C/10); colormap gray; axis xy; axis equal; axis tight;

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104 axis off; %averaging, standard deviation of three repeatability tests N=3; Iavg=(I1+I2+I3)/N; average=Iavg/100*(relconc_A+relconc_B+relconc_C)/N; stddev=Iavg*(((average/Iavg-re lconc_A/100).^2+(average/Iavgrelconc_B/100).^2+(average/Iavgrelconc_C/100).^2)/(N-1)).^0.5; stddev1=stddev./average; stddev_av=100*mean(mean(stddev1(:,1:150))); uncertainty=stddev_av*4.3/N^0.5;%95 percent confidence level %fuel mole fraction at mass sampling location (x/H = 2.0, y/H = 0.3) x=300; y=45; fuelmolefr=mean(mean(average (x-3:x+3,y-3:y+3)))/Iavg*100; fuelmolefr_stddev=mean(mean(s tddev1(x-3:x+3,y-3:y+3)))*100; %filtering filter=ones(3,3)/9; average1=conv2(average,filter); %fuel mole fraction distribution %I_filtered=839;%80psi I_filtered=564;%176psi relconc=100*average1/I_filtered; %plotting

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105 figure(1); Z=relconc; [C,h]=contour(Z,[10 20 30 40 50 60 70 80]); clabel(C,h,'manual'); axis xy; axis equal; axis tight; xlim([3 190]); ylim([3 300]); set(gca,'XTick',[3 75 150 190]); set(gca,'XTickLabel',{'0','0.5','1.0','1.25'}); set(gca,'YTick',75:75:300); set(gca,'YTickLabel',{'0.5','1.0','1.5','2.0'}); xlabel('x / H'); ylabel('y / H'); figure(2); image(average1/10); colormap gray; axis xy; axis equal; axis tight; axis off; figure(3);

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106 image(average/10); colormap gray; axis xy; axis equal; axis tight; axis off; lasercharacterization.m %lasercharacterization.m laser_0min='012006a'; laser_40min='012006b'; laser_85min='012006c'; laser_0min_1='012006d'; %time averaging of a set of images [lasermean_0min]=imbatchread(laser_0min); [lasermean_40min]=imbatchread(laser_40min); [lasermean_85min]=imbatchread(laser_85min); [lasermean_0min_1]=imbatchread(laser_0min_1); %laser sheet profile [m,n]=size(lasermean_0min); for i=1:m, lasermean_0min_1D(i)=lasermean_0min(i ,175)/max(lasermean_0min(:,175)); end; [m,n]=size(lasermean_40min);

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107 for i=1:m, lasermean_40min_1D(i)=lasermean_40min( i,175)/max(lasermean_40min(:,175)); end; [m,n]=size(lasermean_85min); for i=1:m, lasermean_85min_1D(i)=lasermean_85min( i,175)/max(lasermean_85min(:,175)); end; [m,n]=size(lasermean_0min_1); for i=1:m, lasermean_0min_1_1D(i)=lasermean_0m in_1(i,175)/max(lasermean_0min_1(:,175)); end; %average, standard deviation N=3; average=(lasermean_0min_1D+laserm ean_40min_1D+lasermean_85min_1D)/N; stddev=(((average-laserm ean_0min_1D).^2+(averagelasermean_40min_1D).^2+(average-la sermean_85min_1D).^2)/(N-1)).^0.5; stddev1=stddev./average; stddev_av=100*mean(stddev1); %plotting figure(1); plot(lasermean_0min_1D,'r'); hold on; plot(lasermea n_40min_1D,'b');

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108 hold on; plot(lasermea n_85min_1D,'g'); hold on; plot(lasermean_0min_1_1D,'k'); axis xy; xlim([1 320]); ylim([0 1]); xlabel('x / H'); ylabel('I / I max'); set(gca,'XTick',0:80:320); set(gca,'XTickLabel',{'0','0.5','1.0','1.5','2.0'}); legend('t = 0min','t = 40min','t = 85min','t = 0min (repeat)');

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109 LIST OF REFERENCES 1 Curran E. T., Scramjet engines: the first forty years, Journal of Propulsion and Power, Vol. 17 (6), pp. 1138-1148, 2001. 2 Billig F. S., Research on supersonic combustion, Journal of Propulsion and Power, Vol. 9 (4), pp. 499-514, 1993. 3 Marren D., Lewis M. and Maurice L. Q ., Experimentation, test and evaluation requirements for future air breathing hypersonic systems, Journal of Propulsion and Power, Vol. 17 (6), pp. 1361-1365, 2001. 4 Kumar A., Drummond J. P., McClinton C. R. and Hunt J. L., Research in hypersonic airbreathing propulsion at the NASA Langley research center, 15th International Symposium on Airbreathing Engines, ISABE-2001-1007, 2001. 5 Rodriguez C. G., Computational fluid dynami cs analysis of the Central Institute of Aviation Motors/NASA scramjet, Journal of Propulsion and Power, Vol. 19 (4), pp. 547-555, 2003. 6 Ben-Yakar A. and Hanson R. K., Cavit y flame-holders for ignition and flame stabilization in scramjets: an overview, Journal of Propulsion and Power, Vol. 17 (4), pp. 869-877, 2001. 7 Gruber M. R., Baurle R. A., Mathur T. a nd Hsu K. Y., Fundamental studies of cavitybased flameholder concepts for supersonic combustors, Journal of Propulsion and Power, Vol. 17 (1), pp. 146-153, 2001. 8 Ali M. C. and Kurian J., Cavity-bas ed injections into supersonic flow, Journal of Propulsion and Power, Vol. 21 (6), pp. 1130-1132, 2005. 9 Mathur T., Gruber M., Jackson K., Donbar J ., Donaldson W., Jackson T. and Billig F., Supersonic combustion experiments with a cavity-based fuel injector, Journal of Propulsion and Power, Vol. 17 (6), pp. 1305-1312, 2001. 10 Yu K. H., Wilson K. J. and Schadow K. C., Effect of flame-holding cavities on supersonic combustion performance, Journal of Propulsion and Power, Vol. 17 (6), pp. 1287-1295, 2001.

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111 24 Ozawa R. I., Survey of basic data on flam e stabilization and propagation for high speed combustion systems, Airforce Technical Report, AFAPL-TR-70-81, The Marquardt Company, 1971. 25 Huellmantel L. W., Ziemer R. W. and Ca mbel A. B., Stabilization of premixed propane-air flames in recessed ducts, Journal of Jet Propulsion, pp. 31-43, Jan 1957. 26 Baxter M. R. and Lefebvre A. H., Flame stabilization in high-ve locity heterogeneous fuel-air mixtures, Journal of Propulsion and Power, Vol. 8 (6), pp. 1138-1143, 1992. 27 Baxter M. R. and Lefebvre A. H., Weak extinction limits of large-scale flameholders, Journal of Engineering for Gas Turbines and Power Transactions of the ASME, Vol. 114, pp. 776-782, 1992. 28 Niioka T., Terada K., Kobayashi H. and Hasegawa S., Flame stabilization characteristics of strut divided in to two parts in supersonic flow, Journal of Propulsion and Power, Vol. 11 (1), pp. 112-116, 1995. 29 Zakkay V., Sinha R. and Medecki H., Res idence time within a wake recirculation region in an axisymmetr ic supersonic flow, Astronautica Acta, Vol. 16, pp. 201-216, 1971. 30 Driscoll J. F. and Rasmussen C. C., Correlation and analysis of blowout limits of flames in high-speed airflows, Journal of Propulsion and Power, Vol. 21 (6), pp. 10351044, 2005. 31 Rasmussen C. C., Driscoll J. F., Hsu K. Y ., Donbar J. M., Gruber M. R. and Carter C. D., Stability limits of cavity-stabilized flames in supersonic flow, Proceedings of the Combustion Institute, Vol. 30, pp. 2825-2833, 2005. 32 Winterfeld G., Stabilization of hydr ogen-air flames in supersonic flow, Modern Research Topics in Aerospace Propulsion, pp. 37-47, 1991. 33 Rasmussen C. C., Driscoll J. F., Carter C. D. and Hsu K. Y., Characteristics of cavitystabilized flames in a supersonic flow, Journal of Propulsion and Power, Vol. 21 (4), pp. 765-768, 2005. 34 Hsu K. Y., Carter C., Craf ton J., Gruber M., Donbar J., Ma thur T., Schommer D. and Terry W., Fuel distribution about a cav ity flameholder in supersonic flow, AIAA Paper, AIAA-2000-3585, 2000. 35 Gruber M. R., Donbar J. M., Carter C. D. and Hsu K. Y., Mixing and combustion studies using cavity based flameholders in a supersonic flow, Journal of Propulsion and Power, Vol. 20 (5), pp. 769-778, 2004.

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112 36 Uchiumi M., Kobayashi H., Hasegawa S. and Niioka T., Experiments on the flameholding mechanism of a newly de vised strut in supersonic airflow, IUTAM Symposium on Combustion in Supersonic Flows, pp. 135-144, 1997. 37 Zamma Y., Shiba H., Masuya G., Tomioka S ., Hiraiwa T. and Mitani T., Similarity parameters of pre-ignition flowfi elds in a supersonic combustor, AIAA Paper, AIAA1997-2890, 1997. 38 Tomioka S., Hiraiwa T., Mitani T., Zamma Y., Shiba H. and Masuya G., Auto ignition in a supersonic combustor with pe rpendicular injection behind backward-facing step, AIAA Paper, AIAA-1997-2889, 1997. 39 Masuya G., Kudou K., Muraka mi A., Komuro T., Tani K., Kanda T., Wakamatsu Y., Chinzei N., Sayama M., Ohwaki K. and Kimura I., Some governing parameters of plasma torch igniter/flameholder in a scramjet combustor, Journal of Propulsion and Power, Vol. 9 (2), pp. 176-181, 1993. 40 Thayer W. J. and Corlett R. C., Gas dynamic and transport phenomena in the twodimensional jet interaction flowfield, AIAA Journal, Vol. 10 (4), pp. 488-493, 1972. 41 McDaniel J., Fletcher D., Hartfield R. and Hollo S., Staged transverse injection into Mach 2 flow behind a rearwa rd-facing step: a 3-D compre ssible test case for hypersonic combustor code validation, AIAA Paper, AIAA-1991-5071, 1991. 42 Strokin V. and Grachev V., Possible sc heme of flameholding in hydrogen fueled scramjet combustors, Proceeding of 1st International Aerospace Congress, Vol. 1, pp. 630-633, 1997. 43 Ogorodnikov D. A., Vinogradov V. A., Shikhman Y. M. and Strokin V. Y., Russian research on experimental hydr ogen-fueled dual-mode scramj et: conception and preflight tests, Journal of Propulsion and Power, Vol. 17 (5), pp. 1041-1048, 2001. 44 Morrison C. Q., Campbell R. L., Edelman R. B. and Jaul W. K., Hydrocarbon fueled dual mode ramjet/scramjet concept evaluation, Proceedings of the International Society for Air-breathing Engines, ISABE-1997-7053, pp. 348-356, 1997. 45 Kim C. K., Yu S. T. and Zhang Z. C., Cavity flow in scramjet engine by space-time conservation and solution element method, AIAA Journal, Vol. 42 (5), pp. 912-919, 2004. 46 Correa S. M. and Warren R. E., Supers onic sudden-expansion flow with fluid injection: an experimental and computational study, AIAA Paper, AIAA-1989-0389, 1989.

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113 47 Glawe D. D., Donbar J. M., Nejad A. S ., Sekar B., Chen T. H., Samimy M. and Driscoll J. F., Parallel fuel injection from the base of an extended strut into supersonic flow, AIAA Paper, AIAA-1994-0711, 1994. 48 Mitani T., Takahashi M., Tomioka S., Hiraiwa T. and Tani K., Analyses and application of gas sampling to scramjet engine testing, Journal of Propulsion and Power, Vol. 15, No. 4, 1999. 49 Chinzei N., Komuro T., Kudou K., Murakami A., Tani K., Masuya G. and Wakamatsu Y., Effects of injector geometry on scramjet combustor performance, Journal of Propulsion and Power, Vol. 9 (1), pp. 146-152, 1993. 50 Masuya G., Komuro T., Murakami A., Shinozaki N., Nakamura A., Murayama M. and Ohwaki K., Ignition and combustion performance of scramjet combustors with fuel injection struts, Journal of Propulsion and Power, Vol. 11 (2), pp. 301-307, 1995. 51 Ciezki H. K., Scheel F. and Kwan W., Inve stigation of the com bustion process in a scramjet model combustor with a sampling probe system, AIAA Paper, AIAA-20044166, 2004. 52 Rogers R. C., A study of the mixing of hydrogen injected normal to a supersonic airstream, NASA Technical Note, NASA TN D-6114, 1971. 53 Ng W. F., Kwok F. T. and Ninnemann T. A. A concentration probe for the study of mixing in supersonic shear flows, AIAA Paper, AIAA-1989-2459, 1989. 54 Cox S. K., Fuller R. P., Schetz J. A. and Walters R. W., Vortical interactions generated by an injector array to enhance mixing in supersonic flow, AIAA Paper, AIAA-1994-0708, 1994. 55 Hanson R. K., Seitzman J. M. and Paul P. H., Planar laser-fluorescence imaging of combustion gases, Applied Physics B, Vol. 50, pp. 441-454, 1990. 56 Schulz C. and Sick V., Tracer-LIF diagnostics: quantitative measurement of fuel concentration, temperature and fuel-air ratio in practical combustion systems, Progress in Energy and Combustion Science, Vol. 31, pp. 75-121, 2005. 57 Fox J. S., Houwing A. F., Danehy P. M., Gaston M. J., Mudford N. R. and Gai S. L., Mole fraction-sensitive imaging of hypermixing shear layers, Journal of Propulsion and Power, Vol. 17 (2), pp. 284-292, 2001. 58 Hartfield R. J., Abbitt J. D. and McDaniel J. C., Injectant mole fraction imaging in compressible mixing flows using planar laser induced iodine fluorescence, Optics Letters, Vol. 14 (16), pp. 850-852, 1989.

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114 59 Abbitt J. D., Hartfield R. J. and McDaniel J. C., Mole fraction imaging of transverse injection in a ducted supersonic flow, AIAA Journal, Vol. 29 (3), pp. 431-435, 1991. 60 Hollo S. D., McDaniel J. C. and Hartfield R. J., Quantitative investigation of compressible mixing: staged transver se injection into Mach 2 flow, AIAA Journal, Vol. 32 (3), pp. 528-534, 1994. 61 Hartfield R. J., Hollo S. D. and McDaniel J. C., Experimental investigation of a supersonic swept ramp injector using laser induced iodine fluorescence, Journal of Propulsion and Power, Vol. 10 (1), pp. 129-135, 1994. 62 Lozano A., Yip B. and Hanson R. K., Acetone: a tracer for concentration measurements in gaseous flows by planar laser-induced fluorescence, Experiments in Fluids, Vol. 13, pp. 369-376, 1992. 63 Bryant R. A., Donbar J. M. and Driscoll J. F., Acetone LIF for flow visualization at temperatures below 300K, AIAA Paper, AIAA-1997-0156, 1997. 64 Thurber M. C., Grisch F., Kirby B. J., Votsmeier M. and Hanson R. K., Measurements and modeling of acetone laserinduced fluorescence with implications for temperature-imaging diagnostics, Applied Optics, Vol. 37 (21), pp. 4963-4978, 1998. 65 VanLerberghe W. M., Santiago J. C., Dutton J. C. and Lucht R. P., Mixing of a sonic transverse jet injected into a supersonic flow, AIAA Journal, Vol. 38 (3), pp. 470-479, 2000. 66 Hartfield R. J., Hollo S. D. and McDaniel J. C., Pla nar measurement technique for compressible flows using laserinduced iodine fluorescence, AIAA Journal, Vol. 31 (3), pp. 483-490, 1993. 67 OByrne S., Danehy P. M. and Cutler A. D., Dual-pump CARS thermometry and species concentration measurements in a supersonic combustor, AIAA Paper, AIAA2004-0710, 2004. 68 Cutler A. D., Danehy P. M., OByrne S ., Rodriguez C. G. and Drummond J. P., Supersonic combustion experiments for CF D model development and validation, AIAA Paper, AIAA-2004-0266, 2004. 69 Kasal P., Gerlinger P., Walther R., Wo lfersdorf J. and Weigand B., Supersonic combustion: fundamental investigati ons of aerothermodynamic key problems, AIAA Paper, AIAA-2002-5119, 2002. 70 Owens M. and Segal C., Devel opment of a hybrid-fuzzy ai r temperature controller for a supersonic combustion test facility, Experiments in Fluids, Vol. 31, no. 1, pp. 26-33, 2001.

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115 BIOGRAPHICAL SKETCH Amit Thakur was born in 1976, in India. He was raised in Bhilai, a town in central India, known for the steel manufacturing plant lo cated there. In 1994, he enrolled at the Indian Institute of Technology-Kharagpur, co nsidered one of the best engineering colleges in India, and received a Bachelor of Technology in ocean engineering and naval architecture in 1998. Thereafter until 2000, he worked at Zentech in Bombay-India, a company headquartered in Houston-USA and offering engineering c onsultancy services to the offshore-oil drilling and production industry. In 2000, he enrolled at the University of Iowa, Iowa City-USA, and in 2002 he received a Master of Science in mechanical engineering with specialization in experimental fluid mechan ics. In 2003, he enrolled in the PhD program in aerospace engineering at the University of Florida. At the combustion laboratory, he worked on scramjet engines and gained valuable practical experience ranging from setting up, performing and analyzing scientif ic experiments to writing technical papers and delivering presen tations at internati onal conferences. His research interest is in experimental fluid mechanics.


Permanent Link: http://ufdc.ufl.edu/UFE0013100/00001

Material Information

Title: Non-premixed Conditions in the Flameholding Recirculation Region behind a Step in Supersonic Flow
Physical Description: Mixed Material
Copyright Date: 2008

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Source Institution: University of Florida
Holding Location: University of Florida
Rights Management: All rights reserved by the source institution and holding location.
System ID: UFE0013100:00001

Permanent Link: http://ufdc.ufl.edu/UFE0013100/00001

Material Information

Title: Non-premixed Conditions in the Flameholding Recirculation Region behind a Step in Supersonic Flow
Physical Description: Mixed Material
Copyright Date: 2008

Record Information

Source Institution: University of Florida
Holding Location: University of Florida
Rights Management: All rights reserved by the source institution and holding location.
System ID: UFE0013100:00001


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NON-PREMIXED CONDITIONS IN THE FLAMEHOLDING RECIRCULATION
REGION BEHIND A STEP IN SUPERSONIC FLOW















By

AMIT THAKUR


A DISSERTATION PRESENTED TO THE GRADUATE SCHOOL
OF THE UNIVERSITY OF FLORIDA IN PARTIAL FULFILLMENT
OF THE REQUIREMENTS FOR THE DEGREE OF
DOCTOR OF PHILOSOPHY

UNIVERSITY OF FLORIDA


2006

































Copyright 2006

by

AMIT THAKUR















ACKNOWLEDGMENTS

I am grateful to my advisor, Dr. Corin Segal, for guiding me through the research

work and for driving me to finish the Ph.D. program in a relatively short time. Working

at the combustion laboratory helped me learn about scramjet engines, and provided

valuable practical experience ranging from setting up, performing and analyzing

scientific experiments to writing technical papers and delivering presentations at

international conferences. Beyond academics, interaction with fellow lab mates having

diverse backgrounds from different countries made for an enjoyable and educational

experience in itself. Finally, I thank my mother and father for being typical Indian

parents, encouraging their child to pursue higher education.

This work was performed with support from NASA grant NCC3-994, with Claudia

Meyer as the Program Manager.
















TABLE OF CONTENTS



A C K N O W L E D G M E N T S ......... .................................................................................... iii

LIST OF TABLES ........ ........ ................................... ................. ............ vi

L IST O F FIG U R E S .... ...... ...................... ........................ .. ....... .............. vii

ABSTRACT .............. .................. .......... .............. ix

CHAPTER

1. IN TR OD U CTION ............................................... .. ......................... ..

Hypersonic Air-Breathing Vehicle ........ .. ....... .......... ............................ ......... 1
A ir-breathing Engine Classification ........................................ ......... ............... 2
T u rb ojet E n g in e ........................................................................... 2
R am jet E engine ............................................................................... .................
Scramj et Engine .................................. ......................... ........... 4
Flam holding in Supersonic Flow ..................................................... ...............
Cavity Flam eholder ................................. ....................................8
Strut Flameholder .............. ... ......... ......................... 11
Free Shear Layer Fluid Dynamics ....... ......... ............ .... ...... 13
Flam e Stability Lim its ................ .................................. ....... 16
Flame Stability in High-Speed Subsonic Flow ...................................... 16
Non-Premixed Flame Stability in Supersonic Flow......................................19
Species Distribution in Non-Premixed Flameholding Region in Supersonic Flow ...23
E x p erim en ts ............................................. ............................ 2 3
Computations .... ................ ...... .. ...... ... .......................27
Mass Sampling for Species Concentration Measurement in Supersonic Flow ..........27
Optical Diagnostics for Species Concentration Measurement in Supersonic Flow ...29
Planar Laser-Induced Fluorescence (PLIF) .............................. 29
R am an Scattering............................................... 33
Scope of Study .................. ...... .... ............... ........34

2. EXPERIM ENTAL SETUP ................................................. ........ 39

Supersonic W ind Tunnel .............................................................. 39
M ass Sampling and Analysis....................................... ......... 40
H ardw are ....................................................... 4 0









D ata P ro c e ssin g ............................................................................................. 4 3
O ptical D iagnostics.......... ..... ............................................................... ....... .. ...... .. 44
H ardw are ........................................44
Im age A acquisition and Processing ........................................... .....................46

3. RESULTS: NON-REACTING FLOW ........................................... ............... 57

M ass Spectrom etry (M S)......................................... .............................................57
B ase Injection: H elium ......................................................... ............... 57
B ase Injection: A rgon ................................................. ............................. 59
U stream Injection: H elium ........................................ .......................... 61
U pstream Injection: A rgon ............................................................. ... ............ 62
Planar Laser Induced Fluorescence (PLIF)..............................................................63
B ase Injection: H elium ......................................................... ............... 63
B ase Injection: A rgon ............ ...................................................... ........ 64
U stream Injection ..................................................... ..... .......... .. ......... 65
Comparison between MS and PLIF data............................................... ...............65

4. RESULTS: REACTING FLOW ........................................ .......................... 78

B ase Injection: H ydrogen ........................................................................ 78
Upstream Injection: Hydrogen ............................................................................80

5. C O N C L U SIO N S ..................... .... ............................ ........... ...... ... ...... 88

APPENDIX

A. LABVIEW PROGRAM FOR MASS SAMPLING SEQUENCING ........................ 92

B. MATLAB PROGRAM FOR PLIF IMAGE PROCESSING............... .................99

L IST O F R E F E R E N C E S ...................................................................... ..................... 109

BIOGRAPHICAL SKETCH ........................................................................115


















v
















LIST OF TABLES

Table p


2-1. Mass spectrometer calibration factors for various gases............... ...................50

3-1. Base fuel injection: global and local XHe.......................................... ................. 67

3-2. Base fuel injection: global and local X ..................... .............. 67

3-3. Upstream fuel injection: global and local XHe ................................. ................. 67

3-4. Upstream fuel injection: global and local X. ............. ............... 67

4-1. Base fuel injection: global and local OH2 .................. ................................ 82
















LIST OF FIGURES


Figure page

1-1. Comparison between (a) Rocket-powered vehicle [Source: www.isro.org] and
(b) Scramjet-powered hypersonic vehicle [Source: www.nasa.gov].......................36

1-2. Schematic diagram of (a) turbojet engine (b) ramjet engine (c) scramjet engine. ...37

1-3. Various flameholder geometries for supersonic flow: (a) rearward step (b) cavity
(c) strut. ..................................................................................3 8

1-4. Schematic diagram of supersonic flow over a rearward step...............................38

2-1. Description of the test section showing fuel injection and mass sampling
locations: (a) image and (b) schematic diagram .....................................................51

2-2. Mass sampling from the recirculation region behind the step for analysis by the
mass spectrometer: (a) schematic diagram (b) image. ...........................................52

2-3. MS measurements of species mole fraction distribution in the recirculation
region as the solenoid valves switch sequentially from one sampling port to
another under steady experimental conditions (a) non-reacting flow with helium
as the simulant fuel (b) reacting flow with hydrogen combustion.........................53

2-4. Schematic diagram of planar laser-induced fluorescence (PLIF) setup...............54

2-5. Sample PLIF image for (a) background (b) laser sheet. ........................................55

2-6. Temporal variation of laser sheet profile at y/H = 1.1 ........................ ..........56

3-1. Wall pressure distribution for non-reacting flow. Pair = 4.8 atm, Mair = 1.6. The
axial origin is placed at the step. ........................................ .......................... 68

3-2. Base fuel injection: MS measurement of helium mole fraction distribution in the
recirculation region for (a) wall sampling (b) inflow sampling. Pair = 4.8 atm,
M air = 1.6. ..................... ........................................... ........... .69

3-3. Base fuel injection: MS measurement of argon mole fraction distribution in the
recirculation region for (a) wall sampling (b) inflow sampling. Pair = 4.8 atm,
M air = 1.6. ..................... ........................................... ........... .70









3-4. Upstream fuel injection: MS measurement of helium mole fraction distribution
in the recirculation region for (a) wall sampling (b) inflow sampling. Pair = 4.8
atm M air = 1.6 ..................................................... ................. 7 1

3-5. Upstream fuel injection: MS measurement of argon mole fraction distribution in
the recirculation region for (a) wall sampling (b) inflow sampling. Pair = 4.8
atm M air = 1.6 ..................................................... ................. 72

3-6. PLIF measurement for base injection of helium (a) image (b) XHe distribution
(%). POHe = 5.4 atm, Poair = 4.8 atm, M air = 1.6 .............. ..................................... 73

3-7. PLIF measurement for base injection of helium (a) image (b) XHe distribution
(%). POHe = 12.0 atm, Poair = 4.8 atm, Mair = 1.6...................... ...................74

3-8. PLIF measurement for base injection of argon (a) image (b) XA distribution (%).
POAr = 5.4 atm Poair = 4.8 atm M air = 1.6. ..................................... ...............75

3-9. PLIF measurement for base injection of argon (a) image (b) XA distribution (%).
PoAr = 12.0 atm, Poair = 4.8 atm, M air = 1.6. ................................... ............... 76

3-10. Comparison between MS and PLIF data for base fuel injection of (a) helium (b)
arg o n .......................................................................................7 7

4-1. Wall pressure distribution for hydrogen combustion tests. Pair = 4.5 atm, Mair =
1 .6 ...................................................................................... ...8 3

4-2. Base fuel injection: Wall sampling results for (a) hydrogen equivalence ratio
distribution in the recirculation region (b) combustion species mole fraction
distribution. POH2 = 4.5 atm, Poair = 4.5 atm, Mair = 1.6. ........................................84

4-3. Base fuel injection: Wall sampling results for (a) hydrogen equivalence ratio
distribution in the recirculation region (b) combustion species mole fraction
distribution. POH2 = 8.2 atm, Poair = 4.5 atm, Mair = 1.6. ........................................85

4-4. Base fuel injection: Inflow sampling results for (a) hydrogen equivalence ratio
distribution in the recirculation region (b) combustion species mole fraction
distribution. POH2 = 4.5 atm, Poair = 4.5 atm, Mair = 1.6. .........................................86

4-5. Base fuel injection: Inflow sampling results for (a) hydrogen equivalence ratio
distribution in the recirculation region (b) combustion species mole fraction
distribution. POH2 = 8.2 atm, Poair = 4.5 atm, Mair = 1.6. .........................................87















Abstract of Dissertation Presented to the Graduate School
of the University of Florida in Partial Fulfillment of the
Requirements for the Degree of Doctor of Philosophy

NON-PREMIXED CONDITIONS IN THE FLAMEHOLDING RECIRCULATION
REGION BEHIND A STEP IN SUPERSONIC FLOW

By

Amit Thakur

May 2006

Chair: Corin Segal
Major Department: Mechanical and Aerospace Engineering

Flameholding in supersonic flow depends on local conditions in the recirculation

region, and on mass transfer into and out of this region. Large gradients in local gas

composition and temperature exist in the recirculation region. Hence, stability parameter

correlations developed for premixed flames cannot be used to determine blowout stability

limits for non-premixed flames encountered in practical devices. In the present study,

mixture samples were extracted at different locations in the recirculation region and the

shear layer formed behind a rearward-facing step in supersonic flow, and analyzed by

mass spectrometry to determine the species concentration distribution in the region. The

point-wise mass spectrometer measurements were complemented by acetone planar laser-

induced fluorescence (PLIF) measurements to get a planar distribution of fuel mole

fraction in the recirculation region. Non-reacting flow tests and combustion experiments

were performed by varying various fuel related parameters such as injection location,

injection pressure and fuel type. Fuel injection upstream of the step was not effective in









supplying enough fuel to the recirculation region and did not sustain the flame in

combustion experiments. Fuel injection at the step base was effective in sustaining the

flame. For base injection, the local fuel mole fraction in the recirculation region

determined from experiments was an order of magnitude higher than the global fuel mole

fraction based on total moles of air flowing through the test section and total fuel injected

in the test section. This suggests substantial difference in flame stability curve for non-

premixed conditions in the scramjet engine compared to premixed flow. For base

injection, fuel remained in the recirculation region even at higher injection pressure. Due

to slower diffusion rate, the heavier fuel had higher local mole fraction in the

recirculation region compared to lighter fuel for a unit global fuel mole fraction injected

in the test section. Hence fuel molecular weight will affect the non-premixed flame

stability limits in scramjet engine; the heavier fuel will have better fuel-lean and worse

fuel-rich stability limit compared to lighter fuel. This is in addition to the fact that a

lighter fuel such as hydrogen has a much wider flame stability limit than a heavier fuel

such as propane. The data obtained in the study can help develop a stability parameter for

non-premixed flames and validate computational models.














CHAPTER 1
INTRODUCTION

Hypersonic Air-Breathing Vehicle

Hypersonic air-breathing vehicles for launching payloads in space are an area of

active interest in aerospace communities around the world. Such a vehicle could be used

as the first stage of a two-stage-to-orbit launch vehicle with rocket-powered second stage

required to reach orbit. In an optimistic scenario, the hypersonic air-breathing vehicle

could also be used as a single-stage-to-orbit launcher. The existing rocket-based launch

vehicles [Figure 1-1 (a)] carry fuel and oxidizer along with them, and the latter

contributes substantially to the launch vehicle total weight. The oxidizer weight can be as

high as 65% of the total weight while the useful payload is only about 2-3% of the total

weight. This results in high launch costs and limits frequent access to space. An

airbreathing launch vehicle [Figure 1-1 (b)] that propels itself by using oxygen from the

atmosphere instead of carrying it onboard would free up a substantial portion of its

weight that could potentially be utilized to significantly increase the useful payload

weight as a fraction of total vehicle weight and reduce launch costs.

However, there are several challenges in realizing such a hypersonic air-breathing

vehicle. It will fly in the atmosphere at high Mach numbers for a much longer time than a

rocket, and thus will encounter significant drag heating on its airframe. Hence the

airframe should be made of materials that can withstand extremely high temperatures; it

also needs to be actively cooled by cryogenic fuel circulating under the vehicle skin. The

air-breathing engine required to propel the vehicle at such high speeds is a significant









challenge in itself. It is instructive to note that the maximum speed attained by an air-

breathing engine powered aircraft is Mach 3.2; hence no air-breathing engine exists today

for propelling an aircraft at hypersonic speeds.

Air-breathing Engine Classification

Propulsion for an air-breathing launch vehicle would involve combined-cycle

engines that utilize different forms of air-breathing engines most suited for different

stages of the flight envelope. The vehicle may take off from a runway like a conventional

airplane using a turbo-machinery based air-breathing engine. As the flight Mach number

increases in the supersonic regime and the operating limit for turbine-based engine is

reached, the engine shifts to a ramjet mode of operation. At even higher flight Mach

numbers approaching hypersonic speeds, the engine operation shifts to a scramjet mode.

Turbojet Engine

A schematic diagram of turbojet engine is shown in Figure 1-2 (a). Subsonic air

entering the engine inlet is slowed down to low subsonic speed by a diffuser. Its

stagnation pressure is increased by a compressor. Heat is added to the high pressure air

by burning fuel in the combustor and its stagnation temperature is raised further. Fuel and

air are mixed before ignition for uniform heat addition to the airflow; the flame in the

combustor is anchored by a bluff body flameholder. The high pressure, high temperature

air is expanded through the turbine; the work done on the turbine is used to drive the

compressor. The air exiting the turbine still has high temperature; this thermal energy is

converted to kinetic energy by accelerating the air through a converging nozzle. The net

increase in momentum of air passing through the engine produces the thrust needed to

propel the aircraft.









A turbojet engine can also be used to power a supersonic aircraft by a suitably

modified inlet. Oblique shocks at the inlet along with the diffuser slow down supersonic

airflow to subsonic speed suitable for turbojet engine operation. As the airflow speed at

the inlet increases and more compression is achieved by the shock waves, less pressure

rise is achieved at the compressor. Slowing down increasingly higher speed air at the

inlet converts more of kinetic energy to thermal energy and makes it hotter. Increasing its

temperature further by adding heat in the combustor causes air to approach the heat

tolerance limit of the downstream turbine blades. Hence as supersonic speed of the

aircraft increases to about Mach 2.5, the turbojet engine approaches its operational limit.

A ramjet engine is suitable for higher Mach numbers.

Ramjet Engine

A ramjet engine operates only when the vehicle has been accelerated up to

supersonic speed. It has a simple configuration and differs significantly from a turbine-

based engine in the fact that it has no moving parts. A schematic diagram of ramjet

engine is shown in Figure 1-2 (b). A series of oblique shocks at the inlet followed by a

normal shock at the diffuser throat slows the air down to subsonic speed. Sufficient

pressure rise is achieved by the shocks such that the need for a compressor to raise the

airflow pressure and an associated turbine to drive the compressor is eliminated. Heat is

added to the air flowing through the combustor at subsonic speed. The high temperature

achieved due to slowing down of high-speed air helps combustion progress faster in the

engine. A converging-diverging nozzle accelerates the increased enthalpy flow exiting

the combustor and exhausts it out of the engine at supersonic Mach number.









Scramjet Engine

As the flight Mach number increases further in the supersonic regime to about

Mach 5, the operation of a ramjet engine becomes increasingly inefficient. The stagnation

enthalpy of air, comprising kinetic and thermal energy, increases with increasing Mach

number. Shocking such a high energy flow down to subsonic speed converts the kinetic

energy into thermal energy and substantially raises its static temperature. At such high

temperatures, heat addition would only dissociate products into ions and would not add

enthalpy to the flow. Also, a normal shock causes a significant stagnation pressure loss at

high Mach number that is not desirable.

Due to the above mentioned constraints, the high enthalpy airflow cannot be

slowed down significantly by the engine inlet and diffuser, and it enters the combustor at

supersonic speed. So heat is added to the airflow as it flows through the combustor at

supersonic speed, and such an engine is called a supersonic combustion ramjet or

scramjet engine. A schematic diagram of the scramjet engine is shown in Figure 1-2 (c).

The hypersonic Mach number airflow is slowed down to supersonic speed by oblique

shocks at the inlet and then by the diffuser. An isolator is placed before the combustor to

prevent interactions propagating upstream from the combustor to the inlet. In the

combustor, heat is added to air flowing at supersonic Mach number. Typically, the flow

speed entering the combustion chamber of a scramjet engine is about 1/2 1/3 of the

vehicle flight Mach number. The high enthalpy flow exiting the combustor is then

accelerated out of the engine by the converging-diverging nozzle.

There are many challenges in achieving supersonic combustion in a scramjet

engine1l 2, 3 4' 5. Heat needs to be added to supersonic flow in a stable and efficient manner

without causing significant stagnation pressure losses in the process. Efficient fuel-air









mixing and rapid heat release is desirable so that a short combustor length with favorable

thrust to drag ratio can be realized. However, the high-speed air has an extremely short

residence time in the combustor, of the order of a few ms. Hence the fuel gets a very short

time to mix efficiently with air, ignite, undergo complete combustion and add enthalpy to

the incoming flow. Since the airflow entering the combustor is supersonic, the static

pressure and temperature in the combustion chamber is quite low and unfavorable for

rapid chemical reaction. The high-speed airflow is compressible, which has an adverse

effect on the mixing process. Stagnation pressure loss due to heat addition to supersonic

airflow is unavoidable.

If the airflow in the combustor is slowed down to Mach 1 due to heat addition, it

results in thermal choking of the engine. In that case, the disturbances in the combustor

propagate upstream and affect the air intake at the inlet. The static pressure in the

combustor increases significantly and can even lead to blowing up of the engine. Hence

particular care should be taken to avoid thermal choking. This can be achieved by having

a diverging cross-section in downstream part of the combustor, and by limiting the fuel

supply to an acceptable threshold.

The scramjet engine has a lower pressure rise of air entering the inlet, less efficient

heat addition in the combustor, and higher stagnation pressure losses compared to a

turbojet engine. Hence it has a lower thrust to weight ratio than turbojet engine. To

overcome this limitation, bigger scramjet engines are required on a hypersonic aircraft.

This increases the total weight of the aircraft.

Flameholding in Supersonic Flow

Flameholding is an important area of concern in a scramjet engine. A parameter

relevant to flameholding is the Damkohler number (Da) and is defined as the ratio of









flow residence time in the combustor and the fuel reaction time; the reaction time is the

time it takes for fuel-air mixture to mix, ignite and release heat by chemical reaction.


Da = residence (1-1)
reachton

Flameholding is possible only when the residence time available is more than the fuel

reaction time (Da > 1). As mentioned earlier, the flow has a very short residence time in

the combustion chamber and is of the order of only a few ms. In comparison, the

chemical reaction time scales for hydrogen are similar to the scramjet combustor flow

residence time scales, while the chemical reaction time scales for hydrocarbons are much

higher; hence not enough time is available for flameholding in a supersonic combustor.

The short residence time of the flow needs to be increased; hence a solution is to create a

slower, subsonic recirculation region in part of the flow. Flow speed in this region is

favorable to anchoring the flame, and the recirculation flow in this region serves to mix

fuel and air together. The static temperature rise in the recirculation region due to slowing

down of high enthalpy flow reduces the chemical reaction time scale. Hence Da > 1 in

the flameholding recirculation region. The flameholding region thus formed serves as a

reservoir of hot pool of radicals that sustains the flame in the scramjet combustor, and

also as a supplier of radicals helping to propagate combustion in the main supersonic

flow.

The flameholding region discussed above can be created in various ways. In

turbojet engines, a bluff body placed in the combustor main flow is used as a

flameholder. Such an approach is not appropriate in a scramjet combustor since the bluff

body would cause significant stagnation pressure losses in supersonic flow. Various

flameholding geometries suitable for supersonic flow are shown in Figure 1-3. One way









of anchoring the flame in a scramjet engine is to create a recess in the combustor wall in

the form of a rearward step [Figure 1-3 (a)] or a cavity [Figure 1-3 (b)]. A subsonic

recirculation region forms behind the step or in the cavity and acts as the flame anchor.

The main advantage of recess flameholders is that they do not physically obstruct the

supersonic flow and hence avoid stagnation pressure loss. However, they create a very

hot region at the combustor wall that needs active cooling. Since recess flameholders are

located at the wall, they may not able to extend the heat release deep into the main

airflow thus resulting in heat addition only to part of the airflow. Another flameholder is

a slender strut placed in the main flow [Figure 1-3 (c)] and with an appropriate geometry

designed to minimize stagnation pressure losses. The flow separates behind the strut and

forms a recirculation region at the base that anchors the flame. The oblique shocks

generated by the presence of the strut in supersonic flow raise the flow static pressure and

temperature thus assisting in flameholding. However, a drawback is that the strut

experiences very high temperatures since it physically obstructs the high enthalpy

supersonic flow, and hence needs to be actively cooled.

All the geometries discussed above have a common flameholding mechanism,

which is described for a rearward step in Figure 1-4. The approaching boundary layer of

the main airflow separates at the step and forms a shear layer between the supersonic

flow and the subsonic recirculation region. The shear layer is pushed towards the wall

due to the supersonic flow expansion at the step base. An oblique shock is formed at the

shear layer reattachment point. The mass flow rate of air supplied from the main flow

into the recirculation region is governed by the shear layer, as is the supply of hot

combustion radicals from the recirculation region out into the main flow. Fuel is usually









injected in the recirculation region where it mixes with the ambient air through the shear

layer formed at the fuel jet. The flow pattern behind the step is complex in nature with

smaller recirculation zones formed at the covers besides the primary recirculation region.

Further, the figure depicts only the 2-D flow pattern; secondary recirculation regions

exist at the side walls and the flow pattern is 3-D in nature. Hence combustion occurring

in the flameholding region is non-premixed in nature.

Cavity Flameholder

Ben-Yakar and Hanson6 reviewed the flow field characteristics of supersonic flow

over a cavity. Based on the L / D ratio of the cavity, the flow can be characterized as

open or closed. Flow over cavity with L / D < 7-10 is characterized as open flow; the free

shear layer separated at the cavity front wall reattaches at the back wall. For small aspect

ratio cavity with L / D < 2-3, transverse acoustic waves oscillate along the cavity depth.

For larger aspect ratio cavity, the waves oscillate along the cavity length. Cavity flow

with L / D > 10-13 is characterized as closed flow; the free shear layer reattaches at the

cavity floor. Closed flow cavity experiences much higher drag than open flow cavity, the

drag force increases with increasing L / D ratio. The flow residence time in a cavity

increases with cavity depth and the mass entrainment rate increases with cavity length.

The longitudinal oscillations in the cavity are caused due to shear layer

impingement on the back wall. This unsteadiness in flow field is attractive for promoting

fuel-air mixing; however it is undesirable for flameholding. The oscillations can be

passively controlled by angled back wall that prevents reflected acoustic waves. Active

flow control over the cavity can be achieved by upstream injection to enhance the shear

layer growth rate and alter its instability characteristics.









Gruber et al.7 experimentally and computationally evaluated the flow field

properties of various cavity geometries in supersonic flow. Reducing the back wall angle,

defined with respect to horizontal, increased the drag coefficient and shortened the flow

residence time within the cavity. The increase in drag is due to higher pressures acting

over a larger fraction of back wall area. The reduction in residence time is explained by

the structure of recirculating flow within the cavity. Simulations showed that a primary

recirculation region and a secondary embedded vortex exist within the cavity. As the

back wall angle is reduced, the primary recirculation region increases in size and the

secondary recirculation zone is diminished. For such a cavity, the mass exchange takes

place between the primary recirculation zone and the high speed mainstream flow. For a

cavity with higher back wall angle, the mass exchange is slower since part of the

exchange takes place between the low speed recirculation regions. Hence the residence

time reduces with reducing the cavity back wall angle due to higher mass exchange rate.

Ali and Kurian8 used fuel injection into a cavity in supersonic flow as an active

control mechanism for enhancing the air entrainment rate into the cavity due to

interaction between the fuel jet and free shear layer. Fuel injected at different locations

from the cavity floor and front/back wall increased the cavity pressure for all cases. The

pressure fluctuations were suppressed for some of the fuel injection locations.

Mathur et al.9 performed supersonic combustion experiments with cavity

flameholder; fuel was injected upstream of the cavity and from the cavity floor. The

freestream conditions were varied to simulate different stages of hypersonic flight, and

fuel was injected over a range of equivalence ratios. Stable flame anchored in the cavity









and extending out into the main airflow was observed for all experimental conditions.

The engine thrust increased with increasing fuel equivalence ratio.

Yu et al.10 performed supersonic combustion experiments with various cavity

flameholder geometries. Cavity length, back wall angle were varied and cavities in

tandem were used. All cases showed a substantial increase in combustor pressure and

temperature in comparison to the baseline case without cavity. This suggests enhanced

volumetric heat release in the combustor aided by the presence of cavities. Some cavity

configurations such as the inclined cavity and the one with two cavities in tandem

showed a better performance in comparison with other configurations.

Yu et al." performed supersonic combustion experiments with cavity

flameholder; kerosene was injected upstream of the cavity and piloted by hydrogen.

Barbotaged atomization of liquid kerosene using hydrogen increased the combustion

efficiency substantially compared to pure liquid atomization. The combustion efficiency

for fuel injection perpendicular to the airflow was higher than angled injection due to

deeper penetration into the airflow; however it was achieved at the cost of higher

stagnation pressure loss. Combustion performance improved with increasing cavity

depth, which increased the flow residence time, but did not vary much with further

increase beyond a certain depth. Tandem cavities performed better than a single cavity.

Owens et al.12 performed supersonic combustion experiments to examine the

flame stability characteristics of cavity flameholder. Kerosene was injected upstream in

the flow boundary layer and piloted by hydrogen injected in the cavity. Temperature

measurements in the recirculation region indicated that cavities were regions of rich

mixtures and their flame stability was strongly affected by air stagnation temperature via









changes in the local equivalence ratio. At high air stagnation temperature, rich mixtures

existed in the cavity and injection of kerosene resulted in flame extinction, except for

large hydrogen flow rates that enlarged the recirculation region and entrained additional

air from the main flow. At low air stagnation temperature, the flame held even for large

kerosene flow rates.

Davis and Bowersox13 accessed flameholding properties of cavity in supersonic

flow using a simplified, perfectly stirred reactor analysis. Self-ignition of hydrogen was

achieved at a lower temperature than hydrocarbons; hydrogen had shorter ignition delay

time than hydrocarbons at the same temperature. Heat loss from the cavity reduced the

flammability limits. Once the lower residence time t, is calculated using the model, the

cavity depth D required for flameholding can be estimated as D = t, U / 40, where U is

the main airflow velocity.

Strut Flameholder

Brandstetter et al.14 established flameholding in supersonic flow in the recirculation

region between a strut and a cylinder placed downstream of the strut. The cylinder

surface temperature needed to be above a threshold temperature for flameholding to be

sustained after the ignition source was switched off.

Northam et al.15 evaluated the performance of various strut geometries in

supersonic flow. The strut incorporating the deepest step with perpendicular fuel injection

downstream of the step exhibited the best combustion efficiency. The strut configuration

with staged perpendicular fuel injection exhibited the lowest auto-ignition and

flameholding limit. Hence an optimum strut configuration should incorporate a deep step

with staged perpendicular fuel injection for best mixing and flameholding performance.









Lyubar et al.16 decelerated a small fraction of supersonic airflow in a scramjet

combustor to subsonic conditions to provide a stable ignition and flameholding zone.

This was achieved by an injector with its inner surface shaped like a supersonic diffuser,

combustor and nozzle. Hydrogen was injected perpendicular to the decelerated airflow in

the injector and the main airflow. The strong temperature rise caused due to deceleration

of the airflow in the injector resulted in auto-ignition and a sustained supply of radicals to

the main airflow. The normal shock causing auto-ignition in the injector was prevented

from moving upstream during combustion by providing an orifice to allow for pressure

release from inside the injector to the main airflow. The mass flow from the orifice also

enhanced fuel-air mixing in the main airflow.

Gerlinger and Bruggermann17 performed numerical simulations for strut

flameholder in supersonic flow with hydrogen injected in the flow direction at the base of

the strut; the effect of varying the lip thickness at the injector end was investigated. The

mixing efficiency was nearly independent of lip thickness variation. Changes caused

within the shock wave/expansion pattern at the injector exit due to lip thickness variation

had a moderate influence on the stagnation pressure loss.

Tabejamaat et al.18 performed numerical simulations and experiments to

investigate the effect of geometry variation of a strut in supersonic flow. Hydrogen was

injected parallel to the airflow and into the recirculation region formed at the base of the

strut. Increasing the base height increased the recirculation region size, improved the

mixing efficiency and hence increased the maximum combustion temperature in the

recirculation region. Increasing the slit width of injector exit increased the recirculation

region size, but did not appreciably affect the mixing efficiency and maximum









combustion temperature in the recirculation region. The double slit injection changed the

flow field compared to single slit injection; increasing the distance between the two slits

beyond a certain limit changed the flow field substantially and caused flame extinction.

Gruenig et al.19' 20 performed experiments to investigate flame stabilization and

mixing by pylons in supersonic flow. Pylons can be considered as short-length struts.

Mixing efficiency and combustion performance of various pylon geometries were

investigated. The best combustion performance was exhibited by the pylon with fuel

injected inclined to the airflow and through ramps designed to create stream-wise vortices

that promote mixing. Flame stabilization was achieved by oblique shocks generated by

the pylon; the shocks increased the flow temperature and pressure and caused auto-

ignition of fuel-air mixture after they had mixed downstream of the pylon. A wedge

appropriately positioned downstream of the pylon was used to modify the shock structure

and induce auto-ignition at a shorter combustor length. The shock train was in turn

influenced and pushed upstream by heat release in the combustor.

Oblique shocks are often formed within a scramjet combustor. Huh and Driscoll21

investigated the beneficial effects of shock waves on a supersonic jet flame. Shock waves

optimally positioned with respect to the jet flame altered the flame size and significantly

improved the blowout limits. The shock waves enhance mixing by directing the airflow

radially inwards towards the fuel and thus increase the air entrainment rate. They create

an adverse pressure gradient which increases the recirculation region size. They also

improve the chemical reaction rates by increasing the static temperature and pressure.

Free Shear Layer Fluid Dynamics

The free shear layer plays a crucial role in bringing air from the main flow into the

flameholding recirculation region and in transporting combustion radicals out of the









recirculation region into the main airflow. Hence the free shear layer fluid dynamics is

discussed here. Dimotakis22 reviewed experimental data on turbulent free shear layer

growth, mixing and chemical reactions formed between two uniform gas streams. The

local Reynolds number is given by

Re = AU/v > 104 (1-2)
AU = U1 U2


where Ui, U2 are the velocities of the two streams.

The chemical product formation at a station x can be expressed as

, / = (l/ ) (5 ) (S /m, ) (1-3)

where the first factor measures the shear layer growth, the second shows mixing within

the shear layer, and the third indicates the chemical products formed within the

molecularly mixed layer.

.,\iC r layer gi v,\ ii It depends on several flow parameters like velocity ratio (r),

density ratio (s), convective Mach numbers of the two streams (Mci, Mc2), relative mean

density reduction attributed to heat release for combusting flow (q), and pressure

gradient.

r = U2 /U, (1-4)
S = P2 P
A1c = (U1 U)/a1,M c2 = (U, -U2)/a2

q = (Po P)/Po
where Uc is the velocity of the large-scale turbulent structures in the shear layer. The

incompressible shear layer growth rate increases with an increase in density ratio. The

incompressible shear layer growth rate as a function of velocity and density ratio is

expressed as follows.









S(1- r)(l +s /2) (1- sl/2)/(l + s1/2)
-(r,s)= C, [1- ] (1-5)
x 2(1+ s/2r) 1+2.9(l+r)/(1-r)

Compressibility has an adverse effect on shear layer growth. For compressible flow, the

growth rate decreases with increasing convective Mach number as shown in the relation

below.


'[r,s;M2l]
x = f(MA) = (1- fj)e 3M + f,
3 (1-6)
[r,s;M=, = 0]
x
f, = 0.2

Heat release in subsonic flow results in a moderate decrease in shear layer growth.

The outward displacement velocity due to heat release impedes the entrainment process

and offsets the dilatation effect. The density reduction approaches a limiting value with

higher heat release due a substantially unmixed fluid at high Reynolds number. Pressure

gradient results in an increase or decrease in shear layer growth depending on velocity

and density ratio.

Mixing: Enhanced mixing of the two free streams is obtained when shear layer

instability causes a transition from large, two-dimensional vortices to three dimensional,

fully developed, turbulent flow in the shear layer. However, at high Reynolds numbers, a

substantial fraction of fluid in the shear layer is not molecularly mixed.

The shear layer entrains fluid from the two free streams in an asymmetric way. The

molar entrainment ratio En for compressible shear layer is expressed as follows.

M 1 -r 1
EL = [1 + C, f(Me1)] I
MI +r s (1-7)
C, = 0.68


where M1, M2 are molecular weights of the two streams.









The probability density function (PDF) of concentration measurements across the

width of a shear layer shows that away from the boundaries of the two free streams, the

high-speed fluid fraction distribution in the mixed shear layer tends to a most probable

value of E defined as follows.

SEn (1-8)
E +1


Pitz and Daily23 made 3-D velocity measurements in the free shear layer formed

behind a rearward step in subsonic flow. The shear layer growth rate was higher

compared to other measurements for parallel streams; this was attributed to reverse flow

in the recirculation region behind the step. The shear layer growth rate was same for non-

reacting and reacting flow.

Flame Stability Limits

Flame Stability in High-Speed Subsonic Flow

Ozawa24 compiled experimental results on wake stabilized flames in high-speed

combustion systems. The limitation of the compiled results is that they are applicable to

premixed, subsonic flows while the combustion process inside a scramjet engine is non-

premixed and supersonic.

A bluff body placed in a combustion chamber stabilizes the flame in its wake. They

are axisymmetric in shape (cone, hemisphere, disc) and also two-dimensional (V-gutter,

cylinder, flat plate) with varying degree of bluffness. Flame stabilization depends on

aerodynamics of the flow in the wake of the flameholder. The mass entrainment rate from

the main flow into the recirculation region formed behind the body increases and the flow

residence time in the region decreases with an increase in the degree of bluffness.

However, a compromise has to be made between improved flameholding characteristics









and higher pressure losses of a more bluff body. A 2-D body has a higher mass

entrainment rate in its wake and a longer residence time than a corresponding

axisymmetric body, and hence has a better flameholding performance. Combustion

reduces the mass exchange rate and increases the residence time in the recirculation

region.

Various parameters of the inlet flow affect the flame blowout velocity: pressure,

temperature, turbulence, water vapor contamination. An increase in inlet flow pressure

and temperature has a favorable effect on the blowout velocity, with temperature having a

stronger influence. Higher free-stream turbulence intensity increases the mass exchange

rate behind the flameholder and the resulting quenching effect decreases the blowout

velocity. Higher water vapor content of the inlet flow increases the ignition delay time of

the fuel and hence decreases the blowout velocity. A stability parameter was formulated

consisting of inlet flow velocity, pressure, temperature, and the flameholder type. The

flameholder parameter is the maximum width of the recirculation region formed behind

it. A plot of equivalence ratio vs. stability parameter gives the flame stability curve for a

given fuel. The correlation between stability parameter (SP) and equivalence ratio is

shown below.

T P dd
SP = ( ) 5( P )(f X ) = f() (1-9)
300K latm U d
where q!: premixed equivalence ratio

To : free stream stagnation temperature

P : free stream static pressure

U: free stream velocity

do / d: flameholder shape parameter









The mixing intensity in the recirculation region behind 2-D flameholders is about

30% higher than axisymmetric flameholders. However, the highest mixing intensity

achieved in stirred reactors is two to three times higher than bluff body flameholders.

Huellmantel et al.25 performed experiments for stabilizing premixed, subsonic,

propane-air flames using a cavity recess in the combustor wall as the flameholder.

Flameholding performances of various cavity geometries were studied. The flame

stabilization curve was generated for each cavity by obtaining the flow velocity at which

flame blow out occurred as a function of both fuel-lean and fuel-rich equivalence ratios.

A long cavity had a wider flame stability range than a short cavity, and a deep cavity

performed better than a shallow cavity. It indicates the necessity of having sufficient

recirculation volume for achieving good stability and shows that increasing the size of the

flameholder increases its performance. The degree of slope of the downstream end of the

cavity did not affect flame stabilization appreciably.

The cavity flameholder had a superior blowout performance when compared to a

900 V-gutter, which is a standard bluff body flameholder. Also, the cavity caused much

less pressure loss in the main flow since it did not physically obstruct the flow. However,

a cause for concern in using a cavity flameholder is the excessive heating of engine wall

in its vicinity.

Baxter and Lefebvre26 determined the fuel-lean flame blowout limits for high-speed

subsonic afterburner combustor systems by varying a range of parameters such as the V-

shaped flameholder geometry, injector-flameholder spacing, airflow Mach number and

stagnation temperature. The effects of changing some of these parameters were at

variance with homogenous results27. The heterogeneous flame blowout equivalence ratios









were leaner than the homogenous values due to richer local fuel-concentration in the

flameholding region. An approximate analysis was developed to model heterogeneous

effects such as transport and vaporization of liquid fuel droplets from the injection

location to the flameholder, droplet capture and vaporization at the flameholder surface,

and gas entrainment into the recirculation region. The effective equivalence ratio in the

recirculation region determined from the above analysis and the lean blowout

equivalence ratio for homogenous flame obtained from correlation of experimental data27

were in good agreement with each other.

Non-Premixed Flame Stability in Supersonic Flow

Niioka et al.28 performed flame stabilization tests in supersonic flow using a strut

divided stream-wise into two parts, with hydrogen injected in the spacing between them.

For different spacing distances between the strut components (L), flame stability plots

were generated by determining the blowout fuel flow rates as a function of airflow

stagnation temperature. The flameholding performance of the strut system was found to

depend greatly on the distance L. The shock/expansion waves formed around to the struts

were not observed to vary much with variation in L. Hence the variation in flameholding

performance was due to the competition between flow residence time and chemical

reaction time in the recirculation region formed in the intermediate region between the

parts. For short and long distances L, the recirculation region was too fuel rich and fuel

lean respectively resulting in large reaction times (Da = Tresidence / Treaction < 1), and hence

did not give good flame stability. For moderate distances L, air entrainment from

supersonic airflow into the recirculation region was adequate and reasonably well mixed.









The reaction time was less than residence time (Da > 1), hence flame stability was

achieved.

Zakkay et al.29 measured the residence time of a gas introduced in the

recirculation region behind an axisymmetric body in supersonic flow for various

experimental conditions. The molar concentration of the gas in the recirculation region

decayed exponentially with time. The dissipation of gas from the recirculation region was

due to diffusion process. The residence time in the recirculation region was of the order

of a few ms and t U/D 75 for laminar and 40 for turbulent flow, where t residence

time in the recirculation region, U freestream velocity, D bluff body base diameter.

The residence time is lower for turbulent flow than laminar flow due to higher diffusion

rate for turbulent flow compared to laminar flow. The residence time was independent of

the concentration of gas in the recirculation region; higher concentration gas dissipated

faster than lower concentration gas thus resulting in the same residence time for both

cases.

Driscoll and Rasmussen30 performed an analysis to develop a correlation for non-

premixed flame stability limits in supersonic flow. The analysis is based on the idea that

the flame base is sustained in the shear layer and not in the recirculation region. Flame

propagation speed along the stoichiometric contour in the shear layer is matched by the

velocity of the incoming gas. Hot products in the recirculation zone preheat the shear

layer gases and increase the flame propagation speed. Flame blowout is governed by the

imbalance between flame propagation speed and gas velocity. Some additional

parameters governing blowout appeared in the non-premixed flame correlation compared

to the premixed flame correlation: fuel injection location relative to the recirculation









region, fuel injection temperature, and stoichiometric fuel mixture fraction. Also, a global

equivalence ratio appeared in the correlation that is different from the local equivalence

ratio in the flameholding recirculation region. The correlation was applied to previously

measured data covering supersonic and subsonic flows; cavities, bluff bodies and struts

as flameholders; hydrogen, ethylene, methane and propane as fuels; and had an average

uncertainty of 55%.

Rasmussen et al.31 examined the stability of hydrocarbon flames in supersonic flow

using cavity flameholders. The effect of the different parameters on lean and rich flame

blowout limits was investigated: fuel type (ethylene, methane), cavity geometry

(rectangular, aft wall ramp cavity), fuel injection location (cavity floor, cavity aft wall),

and airflow Mach number. The blowout limits showed strong dependence on fuel

injection location. For lean blowout limit, the aft wall ramp injection performed better

than floor injection. This is because ramp injection puts fuel directly into the main

recirculation region formed in the cavity, and hence gets distributed throughout the

cavity. Whereas floor injection puts fuel only in the shear layer, hence fuel is unable to

reach much of the cavity. For rich blowout limit, the floor injection performed better than

aft wall ramp injection. The reason is the same as explained above. For floor injection

much of fuel bypasses the cavity and remains concentrated in the shear layer, while ramp

injection floods the cavity with fuel by injecting directly into the recirculation region.

Hence the fuel injection location relative to the recirculation region is an important

parameter in determining flame stability limits. Ethylene had better flame stability than

methane for lean as well as rich limits, since it has a much shorter ignition delay time and

faster flame propagation speed. The airflow Mach number had little effect on lean









stability limit, however Mach 2 flow showed better flame stability than Mach 3 flow for

rich limit. Increase in Mach number reduces the flow pressure, temperature and time

scale, and its effect on flame stability can be better understood by computer simulation.

Winterfeld32 performed flame blowout experiments in supersonic flow using a

contoured cylindrical flameholder and a cone-cylinder flameholder. A recirculation

region formed behind the flameholder. Hydrogen was injected into the recirculation zone

at different angles relative to the airflow (0). The measured recirculation region size was

bigger for 0 = 00 fuel injection angle than for 0 = 900. This is because more fuel was

injected directly into the recirculation region for = 00 while part of the fuel was blown

into the supersonic airflow for 0 = 900. The flame blowout curve plotting normalized fuel

flow rate as a function of blowout airflow Mach number was also sensitive to the fuel

injection angle. For fuel rich limit, better flame stability was achieved for higher injection

angles. It is again explained by the fact that fuel injection at low angle 0 floods the

recirculation with fuel, while at high angle 0 part of fuel escapes into the supersonic

airflow.

Rasmussen et al.33 investigated the effect of fuel injection pressure on flame

location within a cavity in supersonic flow. Fuel was injected from the cavity floor and

from the cavity aft wall; the flame chemiluminescence was captured using a digital

camera. For moderate fuel flow rates with both injection locations, the flame was

anchored in the shear layer and also extended to the recirculation region. For lean and

rich blowout with floor injection, the flame structure was similar to the moderate fueling

case. For lean and rich blowout with aft wall injection, the flame structure showed a

marked departure from moderate fueling case; the flame was not anchored in the shear









layer and was restricted to a region close to the aft wall. The change in flame stabilization

mechanism for aft injection may be attributed to a change in the cavity flow field with

heat release.

Species Distribution in Non-Premixed Flameholding Region in Supersonic Flow

Experiments

Hsu et al.34 used Raman scattering to make fuel distribution measurements inside a

cavity in Mach 2 non-reacting flow. Ethylene was injected at a low angle upstream of the

cavity. The effect of fuel injection pressure, cavity size, and imposed back-pressure on

fuel transport in the cavity was studied. Fuel-rich pockets were observed in the cavity.

For upstream fuel injection, mass transport through the shear layer and its interaction

with the cavity aft wall controls the amount of fuel entering into the cavity. The amount

of fuel entering the cavity decreased as the fuel pressure was increased from moderate to

high values. Higher fuel injection pressure with an increased jet momentum penetrates

into the main airflow and less fuel gets entrained into the shear layer and reaches the

cavity. An increase in cavity size captured more fuel at the cavity back wall, but the drag

penalty also increased. Increase in back pressure, which simulates combustion conditions,

caused the boundary layer upstream of the cavity to separate thus changing the shear

layer interaction with the cavity. The fuel jet penetrated into the main flow due to reduced

momentum of the boundary layer and less of it reached the cavity.

As a follow up of the experiments described above, Gruber et al.35 examined the

effect of fuel injection location, fuel flow rate, and induced back-pressure on the cavity

flameholder performance in supersonic flow. NO-PLIF was used to visualize fuel

distribution in the cavity for non-reacting flow, and OH-PLIF was used in combustion

experiments. From non-reacting flow PLIF visualizations, it was concluded that upstream









fuel injection with passive entrainment of fuel from main airflow into the cavity is less

desirable than direct fuel injection into the cavity. Unlike upstream fuel injection, direct

injection from the cavity aft ramp into the main recirculation region provided a spatially

uniform fuel-air mixture in the cavity. Also, the cavity fuel distribution for direct

injection remained relatively insensitive to changes in the main airflow as simulated by

induced back-pressure. In combustion experiments, as the fuel flow rate from cavity aft

ramp was increased, the cavity was flooded with fuel and adversely affected combustion

in the region. However, when a shock train was imposed for the high fuel flow rate, it

significantly improved cavity combustion by causing the cavity shear layer to separate

which effectively increased air entrainment in the cavity. The shock train could also

enhance fuel-air mixing in the cavity.

Uchiumi et al.36 followed up on the investigation by Niioka at al.28 discussed earlier

and conducted experiments to improve the flameholding performance of a strut divided

into two parts in supersonic flow. Previous non-reacting flow measurements of local

equivalence ratio along the distance between the strut parts showed that for short distance

L, the intervening region was largely fuel rich in composition. Such a mixture required a

long chemical reaction time compared to the short residence time available, and hence

flameholding could not be established. As the distance L was increased, the measured

local equivalence ratios were reduced since more air from the main flow was entrained

into the recirculation region. However, for a certain range of moderate distance L, the

chemical reaction time was more than the flow residence time and hence flameholding

could not be established. Modifications in the fuel-injecting strut were made in the

present experiment. For short distance L between the struts, slits incorporated in the









injection strut brought air from main flow into the recirculation region and diluted its fuel

rich composition. For moderate distance L, a recess was provided around the fuel injector

to prevent excess air from entering the intervening space. The strut modifications resulted

in Da > 1 and hence flameholding was established in both cases.

Zamma et al.37 measured pressure and gas composition for non-reacting supersonic

flow over a step with fuel injected downstream of the step and transverse to the airflow.

Species in the recirculation region behind the step were extracted at the wall and analyzed

by gas chromatography. Fuel was entrained from the jet into the recirculation region

behind the step. The fuel volume fraction in the recirculation region decreased with

increasing fuel-air dynamic pressure ratio. Higher fuel volume fraction was observed in

the recirculation region for lighter fuel compared to heavier fuel injected at the same

dynamic pressure ratio. The gas composition measurements were used to estimate the

ignition characteristics of the flameholder38' 39

Thayer and Corlett40 measured pressure, temperature and gas composition in the

separated recirculation region upstream of a fuel jet injected transverse to non-reacting

supersonic airflow. The species extracted at the wall were analyzed by gas

chromatography. A large part of the recirculation region had a nearly uniform fuel

distribution; the region was fuel-rich in composition. The fuel concentration in the

recirculation region decreased with decreasing fuel mass flow rate. It was estimated that

approximately 5 % of the injected fuel was entrained in the recirculation region.

McDaniel et al.41 made 3-D measurements of flow variables for a flow field with

staged transverse fuel injection downstream of a step in non-reacting supersonic flow.

Pressure, temperature, velocity and fuel mole fraction were measured using LIF. Fuel









was entrained into the recirculation region behind the step from the jet closer to the step

base; fuel concentration close to stoichiometry was observed in the recirculation region.

Strokin and Grachev42 obtained experimental data on ignition and flameholding in

supersonic flow using a cavity flameholder for a variety of experimental conditions; the

results were also reported by Ogorodnikov et al.43. The data was correlated to obtain a

flame stability curve using the air/fuel equivalence ratio in the cavity recirculation zone

and a flameholding parameter based on the airflow velocity, airflow stagnation

temperature, recirculation region static pressure and cavity length. The equivalence ratio

in the recirculation region was empirically estimated using air and fuel stagnation

temperature, recirculation region static temperature, approaching airflow boundary layer

thickness, spacing between fuel injection holes, and overall air/fuel equivalence ratio.

Morrison et al.44 performed experimental and analytical studies on a dual-mode

propane-fueled ramjet/scramjet combustor to determine the conditions necessary to

establish flameholding in the engine. The effects of non-premixed fuel-air mixture and

combustion-induced shock train on flameholding were investigated. Hydrogen was

injected near the exit to raise the back-pressure and establish a shock train that resulted in

subsonic flow over the step flameholder and simulated ramjet conditions. Propane was

injected at a low velocity at the step base to prevent it from penetrating through the

recirculation region. The air entrainment rate into the recirculation region was estimated

analytically from the step base pressure measurements. Premixed conditions were

assumed to exist in the recirculation region and the local equivalence ratio in the region

was calculated from the known fuel flow rate and estimated air entrainment rate. Lean

blowout at a local equivalence ratio close to unity indicated poor fuel-air mixing in the









recirculation region; the mixing efficiency was estimated as 30%. The experiments were

repeated for different levels of back-pressure to determine both fuel-lean and fuel-rich

blowout points and construct a flame stability curve. The curve was similar in shape to

the Ozawa stability loop for premixed subsonic flow24, but it had reduced blowout limits.

Computations

Kim et al.45 reported 2-D numerical simulation of fuel injection over a cavity in

non-reacting supersonic flow. When fuel was injected upstream of the cavity, it

drastically suppressed flow oscillations in the incoming boundary layer leading to a more

stable free shear layer over the cavity. For downstream injection, the interaction between

flow oscillations near the aft cavity wall and the injected jet amplified the cavity pressure

oscillations as compared to the case without injection. Fuel mass fraction contours

showed that large vortices occurred that pulled the fuel upstream into the cavity.

Correa and Warren46 performed 2-D numerical simulations and experiments for

non-reacting supersonic flow over a backward-facing step. Fuel was injected downstream

of the step and transverse to the airflow. The computed fuel mixture fraction showed that

the fuel was confined to the wake of the step. In reality, a 3-D horse shoe vortex forms

around the fuel jet and it penetrates further into the main airflow.

Glawe et al.47 performed 2-D numerical simulations for helium injected at the base

of a strut in non-reacting supersonic flow. CFD contours showed a high helium mole

fraction distribution in the recirculation region formed at the base of the strut.

Mass Sampling for Species Concentration Measurement in Supersonic Flow

Mitani et al.48 performed experiments and computations to analyze and select a

suitable mass sampling probe for a hydrogen supersonic combustor. The extracted

samples were analyzed for species composition using gas chromatography. Different









types of gas sampling probes were examined: a reaction-oriented probe having a uniform

cross-section with no water cooling; a reaction-freezing oriented water cooled probe with

a short cross-section expanding to a larger cross-section to facilitate expansion cooling of

the sampled gas; and a static pressure-type probe to grasp the external flow across the

boundary layer on probe surface. The effectiveness of freezing-oriented probe was

examined by comparison with reaction-oriented probe. Combustion efficiency as a

function of stagnation temperature To was estimated from gas sampling by each of the

probes. The reaction-oriented probe allowed reactions to continue ahead of and inside it

and showed complete combustion for To > 910 K. It gave high but false values for

combustion efficiency. The freezing-oriented probe indicated that combustion was

initiated when To > 1200 K and gave lower combustion efficiency, which were in line

with static-type probe measurements. The validated gas sampling was applied to scramjet

testing under flight Mach numbers up to 8. They clarified distortions of air and hydrogen

in the swept-back, side-compression type engines.

Chinzei et al.49, Masuya et al.50, Ciezki et al.51, Rogers52 used mass sampling and

subsequent analysis by gas chromatography to determine species distribution at different

cross-sections in a supersonic combustor for non-reacting and reacting flow. The mixing

effectiveness of various fuel injection configurations such as step injection, strut

injection, and wall injection normal to the airflow were examined.

Ng et al.53 developed an aspirating hot-film mass sampling probe and used it to

measure local gas composition in a helium-air supersonic shear layer. The probe

consisted of a diverging section followed by a constant area section and a choked orifice,

and was connected to a vacuum pump at the back. The configuration allowed the sampled









flow to expand and form a normal shock in the diverging section of the probe, thus

avoiding a normal shock and flow spillage at the probe tip. The hot-film sensor was

placed in the constant area section of the probe and was connected to an anemometer.

The voltage response of the anemometer is a function of the sampled gas composition,

total pressure and total temperature experienced by the sensor. The total pressure and

temperature at the sensor were measured separately, which allowed for the determination

of sampled gas composition using calibration curves. Cox et al.54 used the probe to study

mixing efficiency of an aerodynamic ramp in supersonic flow.

Optical Diagnostics for Species Concentration Measurement in Supersonic Flow

Planar Laser-Induced Fluorescence (PLIF)

Hanson et al.55, Schulz and Sick56 provided an overview of planar laser-induced

fluorescence (PLIF) as a diagnostics tool for measuring species concentration,

temperature, pressure and velocity in the flow field. The fluorescence signal Sf can be

expressed as follows.


E XP A
Sf = optics /dV, kTabs
hc IA kT A+Q


(1-10)


where optics

E

he /

dVc

X

Uabs

A

Q


efficiency of the collection optics

laser fluence (J/cm2)

energy of a photon

collection volume (cm3)

mole fraction of absorbing molecule

molecular absorption cross section (cm2)

spontaneous emission rate (s-1)

collisional quenching rate (s-1)











The ratio A / (A + Q) gives the fluorescence yield p, which is the fraction of absorbed

photons re-emitted as fluorescence photons. p depends on composition, temperature and

pressure of the gas mixture, and excitation wavelength. oabs depends on gas temperature

and excitation wavelength. Hence the fluorescence signal is a function of species mole

fraction, temperature, pressure and excitation wavelength. Mole fraction imaging is

straightforward for a flow with relatively constant temperature and pressure. For a

relatively constant pressure flow with temperature variation, the mole fraction can be

determined by selecting a suitable excitation wavelength that minimizes the signal's

overall temperature dependence in the temperature range of experiment. Another

approach is to take two almost simultaneous images with excitation at two different

wavelengths. The ratio of the two image signals is a function of temperature only. The

flow field temperature variation thus obtained can be used to determine the mole fraction

distribution from either of the images.

Fox et al.57 used NO PLIF to determine the fuel mole fraction distribution and

evaluate the mixing performance of various fuel injectors in supersonic flow. The

temperature dependence of the signal was minimized by tuning the laser to excitation

wavelengths for which fluorescence is relatively independent of temperature. The

pressure dependence of the signal was minimized by canceling the implicit pressure

dependence of fluorescence yield with explicit pressure dependence of number density;

this was achieved by increasing the quenching cross section to an order of magnitude

higher than the fluorescence cross section.









Hartfield et al.58, Abbitt et al.59, Hollo et al.60, Hartfield et al.61 used iodine PLIF to

measure injectant mole fraction distribution in non-reacting supersonic combustor and

evaluate mixing rates for different injection configurations. The fluorescence intensity

was a function of injectant concentration as well as the local thermodynamic pressure and

temperature. The ratio of the fluorescence signal collected with only the injectant jets

seeded to the signal collected with both the jets and the main flow seeded was taken to

cancel the thermodynamic dependence of fluorescence, and the injectant mole fraction

was obtained as follows.

XN jet = 'e" (1-11)
injectant C (11)
total f-total

where X: mole fraction, N: number density, Sf: signal intensity.

Lozano et al.62 explored the use of acetone (CH3-CO-CH3) as a suitable tracer for

PLIF concentration measurements in gaseous flows. Acetone has a fairly high vapor

pressure (180 torr) at room temperature (293 K) and hence allows a good seeding density,

which can be increased further by increasing the temperature. It absorbs over a broadband

of wavelengths (225-320 nm) with a maximum between 270-280 nm (absorption cross-

section o = 4.7 x 10-20 cm2). The fluorescence emission is broadband in blue (350-550

nm) with peaks at 445 and 480 nm, quantum efficiency 4 = 0.2 %, and a short lifetime of

less than 4 ns. The fluorescence quantum yield is dominated by intersystem crossing;

collision quenching is negligible and hence the yield is independent of local gas

composition. For the experimental conditions investigated, the signal showed no

dependence on temperature and hence was only sensitive to the species concentration.

Acetone phosphoresces in the blue, with a similar emission spectrum as fluorescence









(350-600 nm) and a long decay time (200 |ts for 313 nm excitation). However, it is

greatly quenched by trace amounts of oxygen and also depends on temperature.

Bryant et al.63 conducted experiments to determine the temperature and pressure

dependence of acetone LIF signal. The pressure and temperature range investigated were

0.1 1.0 atm and 240 300 K respectively, conditions akin to that in a supersonic wind

tunnel. For laser excitation wavelength of 266 nm, the LIF signal showed no dependence

on pressure and varied by 5% over the temperature range. Hence the signal can be used to

directly obtain species concentration in an acetone-seeded flow.

Thurber et al.64 investigated the temperature and excitation wavelength dependence

of acetone LIF signal; the temperature range was 300 1000 K and the wavelength range

was 248 320 nm. The signal varied with temperature and wavelength for the range

investigated. At 266 nm, the signal per molecule was constant for 300 350 K and

decreased for higher temperatures.

VanLerberghe et al.65 used acetone PLIF to investigate mixing of an under-

expanded sonic jet injected transversely in a supersonic crossflow. Instantaneous, mean,

standard deviation images and image intensity probability density functions (PDF) were

used to study the mixing process. For the given flow conditions, PLIF signal intensity

was proportional to the acetone mixture fraction; temperature and pressure were

estimated to not play a significant role.

Hartfield et al.66 measured pressure, temperature and velocity in supersonic flow

using iodine PLIF. The uncertainty in measurements due to possible condensation of

iodine in the low temperature supersonic airflow was investigated. Minor condensation

was observed on the test section walls. The signal intensity from the condensed droplets









would be much higher than that from vapor. However, the measured signals were

comparable with the calculated theoretical values. Hence no direct evidence of significant

iodine condensation in the airflow was observed.

Raman Scattering

O'Byrne et al.67 used dual-pump coherent anti-Stokes Raman spectroscopy (CARS)

to measure temperature and mole fraction distribution of N2, 02 and H2 in a supersonic

combustor. A parallel CFD study was done by Cutler et al.68. CARS has the advantage of

producing a coherent signal beam in a particular direction. This increases the signal-to-

noise ratio, and also allows measurements where optical access to the flow is limited.

Three lasers were used in the experiments. The 532 nm beam output from a Nd:YAG

laser was split three ways. The first beam was used as the green pump beam for N2

CARS. The second beam was used to pump a dye laser operating at 554 nm that provided

the yellow pump beam for 02 CARS. The third beam was used to pump another dye laser

operating at 607 nm that provided the red stokes beam for both N2 and 02 CARS. The

frequency difference between blue pump beam and red stokes beam equals a vibration

Raman shift of N2, while the frequency difference between yellow pump beam and red

stokes beam resonates with a vibration Raman shift of 02. The resulting blue CARS

spectrum contains both N2 and 02 spectra. Coincidentally, several pure-rotational Raman

transitions of H2 are also present in this spectral region. The relative intensities of N2, 02

and H2 spectra provided a measure of mole fractions of these species.

Kasal et al.69 evaluated mixing of hydrogen injected from planar and lobed struts in

supersonic non-reacting flow. ID Raman spectroscopy was used to measure species

composition at different cross-sections along the combustor. A 532 nm, 400mJ Nd:YAG









laser was used as the exciting source. The resulting Raman spectrum was used to obtain

mole fractions of H2, 02 and N2 in the flow.

Scope of Study

The aim of this study is to investigate the non-premixed conditions in the

recirculation region formed behind a rearward step in supersonic flow; such a study is of

practical importance in understanding the flameholding mechanism in a scramjet engine

combustor. The step is selected as the flameholder geometry since it is the simplest and

most widely used configuration in a supersonic combustor. The literature review section

discussed earlier shows that only a few investigations have been conducted to determine

the local species distribution in the flameholder recirculation region for non-reacting

supersonic flow. As per the author's knowledge, no such data has been reported for

combustion tests.

This study creates a database of local species mole fraction and fuel equivalence

ratio distribution in the recirculation region formed behind a step in supersonic flow.

Such a database also helps in providing useful input data for complementary

computational fluid dynamics (CFD) efforts. Both non-reacting and combustion cases

were investigated. The airflow parameters such as Mach number, stagnation pressure and

stagnation temperature were held constant. For a given flow condition, the influence of

the following fuel related parameters on the species distribution and combustion in the

flameholding recirculation region were investigated:

* Fuel injection location relative to the recirculation region

S Fuel injection pressure

* Fuel type






35


The tools used in the investigation were a combination of mass sampling and

composition analysis using mass spectrometry (MS) and acetone planar laser induced

fluorescence (PLIF) for visualizing species concentration distribution in non-reacting

flow.





















































Figure 1-1. Comparison between (a) Rocket-powered vehicle [Source: www.isro.org]
and (b) Scramjet-powered hypersonic vehicle [Source: www.nasa.gov].












Combustor
Inlet
Nozzle




M< 1



Compressor Turbine


(a)


Oblique ,
shock ,



M>1



Inlet Combustor Nozzle
M<1

(b)



Oblique ,




M>1
shock -..------







Inlet Isolator Combustor
Nozzle
M>1


(c)



Figure 1-2. Schematic diagram of (a) turbojet engine (b) ramjet engine (c) scramjet
engine.










M>1






(a)


M>1





(b)


M> 1


(c)


Figure 1-3. Various flameholder geometries for supersonic flow: (a) rearward step (b)
cavity (c) strut.


F ree~rea
M> 1

Approaching boundary
layer


I.L. Shear Layer


WIse fuel
direction


Figure 1-4.


Fuel shear layers Prirary recirculation

Schematic diagram of supersonic flow over a rearward step.














CHAPTER 2
EXPERIMENTAL SETUP

Supersonic Wind Tunnel

The supersonic wind tunnel facility used in the experiments, and shown in Figure

2-1 (a), provides direct connect tests with a variable combustion chamber entrance Mach

number of 1.6 3.6 and stagnation temperatures corresponding to Mach 3.0 4.8 flight.

The flight Mach numbers correspond to the transition phase from ramjet to scramjet

engine. The wind tunnel is a continuously operating facility using a vitiated heater based

on hydrogen combustion with oxygen replenishment, electronically controlled by a fuzzy

logic controller70 to maintain a constant 0.21 oxygen mole fraction at all conditions, and

to maintain at the heater exit the constant stagnation temperature profile required for the

experiment. A bell mouth with four-side contraction leads to the supersonic nozzle with

compression on two sides and interchangeable nozzle blocks that cover the range of

Mach 1.6 3.6. All the experiments discussed here were performed with combustion

chamber entrance Mach 1.6 and cold air (Toair = 300 K). A constant area isolator is placed

between the nozzles and the combustor section to protect the nozzle from upstream

pressure rise due to combustion in the test section. Optical access is available to the

isolator's flow from three sides. The isolator cross-section is 2.5 x 2.5 cm2 upstream of a

rectangular, rearward facing step having step height H = 12.7 mm, and follows with a

constant cross-section area test section 26H in length. The test section is symmetric and

has the option of optical access through covering windows. It was water-cooled for

combustion tests.









Fuel was injected transverse to the airflow and into the recirculation region 0.2H

from the base of the step, or transverse to the airflow and at a distance 4H upstream of the

step; a schematic diagram is shown in Figure 2-1 (b). For base injection, five 0.5-mm dia.

holes equally spaced on each side of the test section wall were used. For upstream

injection, two 1.0-mm dia. holes equally spaced on each side of the isolator wall were

used. Helium having molecular weight close to hydrogen, and argon having molecular

weight close to propane, were injected as simulated fuel in non-reacting flow tests.

Hydrogen was injected in combustion tests.

A LabView program and associated National Instruments hardware was used to

monitor experimental conditions such as airflow stagnation pressure, temperature and

Mach number at the nozzle exit, fuel stagnation pressure and temperature, test section

wall static pressure and temperature distribution along the airflow direction. Mass

spectrometry (MS) and planar laser induced fluorescence (PLIF) were used to determine

the species mole fraction distribution in the recirculation region as described below.

Mass Sampling and Analysis

Hardware

The physical location of mass sampling ports in the recirculation region behind

the step is shown in Figure 2-1 (b). The coordinate system is also shown in the figure.

The test section window wall covering the step has five mass sampling ports in the

recirculation region along the axial x-direction equally spaced from x/H = 0.5 to 3.5 and

along y/H = 0.3. These ports are 0.6 mm inner diameter steel tubes that end at the test

section window wall and do not physically intrude into the recirculation region (z = 0). In

separate tests, other tubes are inserted from the window wall to verify the two-

dimensionality of species distribution in the recirculation region. For non-reacting flow









tests, three stainless steel tubes are placed at x/H = 2.0, y/H = 0.3, the location of port #3

in wall sampling; they penetrate into the test section to sample species at three different

depths, equally spaced in the inflow z-direction from z/W = 0.33 to 1.0. Here W = 12.7

mm is the test section half-width. For combustion tests, five 0.8 mm inner diameter

ceramic tubes are inserted to z/W = 0.5 from the wall at the same axial locations as the

wall sampling ports. Exposure to high temperatures of hydrogen combustion flow field in

the flameholding region can cause metals to melt or oxidize. Hence the inflow sampling

tubes for combustion tests were made of ceramics to withstand the high temperatures;

inspection of these tubes after experiments showed that they remained intact during the

tests.

A schematic diagram of mass sampling from the recirculation region and

subsequent, real-time analysis by a mass spectrometer is shown in Figure 2-2 (a). In

combustion tests, the extracted species passed through a water-cooled jacket on their way

to the mass spectrometer to quench the reactions and freeze the species composition

coming out of the combustion chamber. The jacket was supplied with cold water coming

out of a chiller at 283 K. The temperature drop of the sampled mixture while passing

through the cooling jacket resulted in condensation of water vapor and much of it could

not reach the mass spectrometer. Hence the corrected XH20 was deduced from the oxygen

deficit in the product mixture.

The sampling tubes coming out of the recirculation region were connected to a

manifold having six 0.6-mm diameter input tubes and a single 1.8-mm diameter outlet

tube connected to the mass spectrometer. The input of species to the manifold was

regulated by a series of computer-controlled miniature solenoid valves that supply gas









mixture from one sampling port at a time for analysis. Sampling from each port was

preceded by injection of nitrogen in the manifold to purge the line and flush the species

from the previous port, hence preventing mixing of samples from two adjacent ports. A

LabView program and associated National Instruments hardware was used to send analog

signals to operate the miniature solenoid valves in the desired sequence. The sampling

time at each port and the purge time before each sampling were input to the software. The

software is attached in Appendix A. Due to small volume of the sampling system, only a

small mass of species was extracted from the recirculation region, which in turn

facilitated quenching of the reacting species and also helped the mass spectrometer attain

steady state measurements almost instantaneously while switching from one port to

another.

The species were analyzed by Stanford Research Systems RGA-300 mass

spectrometer, shown in Figure 2-2 (b), that uses electron ionization to ionize the sampled

gas, RF quadrupole filter to sort species according to their mass-to-charge ratio, and

Faraday cup to detect ion currents. The instrument is controlled and operated by software

and associated electronics. The ionizer, filter and detector are enclosed in a clean vacuum

chamber and require an operating pressure range of 10-4 torr (1.3 x 10-7 atm) to ultra high

vacuum. Such low pressure is attained in two stages. In first stage, a rotary pump brings

the inlet pressure down to about 60 mtorr (7.9 x 10-5 atm). In second stage, a diffusion

pump and a rotary pump operate in series to bring the pressure further down to vacuum

conditions. The spectrometer can detect species up to a mass-to-charge ratio of 300 and

has a resolution of 0.5 AMU @ 10% peak height. The sensitivity factor of the instrument,

defined as the signal detected per unit partial pressure of a given species (Amp / torr),









varies for different gases. Hence calibration of the instrument was performed for the

following gases: helium, nitrogen, oxygen and argon. The sensitivity factor of nitrogen is

used as the baseline and sensitivity factors of other gases are normalized with this

baseline. The relative sensitivity factor (Sgas / SN2), also referred as the calibration factor,

is shown in Table 2-1 for different gases.

Data Processing

The composition of gas in the recirculation region was analyzed by the mass

spectrometer in partial pressure of species vs. time mode. The species scanned were

nitrogen (m/z=28,14), oxygen (m/z=32,16), helium (m/z=4), argon (m/z=40,20) for non-

reacting experiments, and nitrogen (m/z=28,14), oxygen (m/z=32,16), hydrogen

(m/z=2,1), water (m/z=18,17) for combustion experiments. Sampling was done

sequentially for 5 sec at the purge port and for 20 sec at each of the sampling ports. The

spectrometer had a fast response time of about 2-3 sec to the change in composition of

gas while switching from one port to another. The local mole fraction of a given species

in the sample was determined from the partial pressures of all the component species

recorded by the mass spectrometer. The time-averaged fuel mole fraction at a port was

obtained by averaging the mole fractions obtained over the sampling time period. The

species mole fractions were corrected using calibration factors for individual gases. The

background level of argon in the incoming airflow for non-reacting experiments, and the

background levels of hydrogen and water for combustion experiments were subtracted to

determine the actual mole fractions of these species. Figure 2-3 shows the species mole

fraction distribution in the recirculation region as the solenoid valves switch sequentially

from one sampling port to another under steady experimental conditions; a non-reacting









flow case with helium as the simulant fuel is shown in Figure 2-3 (a) and a reacting flow

case with hydrogen combustion is shown in Figure 2-3 (b).

The global mole fraction/equivalence ratio of fuel was determined from the total

mole of fuel injected and the total mole of air traveling through the test section. Both

local and global mole fractions/equivalence ratios are indicated in the results.

Optical Diagnostics

Hardware

Since mass spectrometry was limited to point-wise measurements in the flow, it

was complemented by acetone planar laser induced fluorescence (PLIF) to obtain fuel

distribution in a 2D plane in the recirculation region for non-reacting flow. The plane of

measurement was along z/W = 0.9. The PLIF measurements overlap with MS data at the

point x/H = 2.0, y/H = 0.3, z/W = 0.9; Figure 2-1 (b) shows the laser sheet in the test

section and the common point between MS and PLIF where measurements from the two

techniques are compared in Chapter 3.

A schematic diagram of PLIF setup is shown in Figure 2-4. The test section used in

PLIF measurements had step on only one side, unlike on both sides for the test section

used in mass spectrometer measurements. However, since the airflow arriving at the step

base is supersonic, the flow field in the recirculation region is the same for both test

sections. The test section side wall window was made of glass for laser sheet delivery in

the recirculation region. The test section front windows next to the step were also made

of glass to provide visual access of the measurement plane via a camera. Before being

injected in the test section, the fuel was seeded with acetone vapor by bubbling it into a

chamber partially filled with liquid acetone at room temperature (295 K).









A pulsed Nd:YAG Spectra-Physics laser was used for LIF excitation of acetone.

The laser pulses at a frequency of 10 Hz and can have wavelength outputs of 1064, 532,

355 and 266 nm. The fourth harmonic 266 nm beam was used for acetone PLIF. At 266

nm wavelength, the laser pulse had 70 mJ energy and a pulse width of 4-5 ns. The beam

diameter coming out of the laser was 7 mm. It was converted into a sheet of light and

delivered in the test section using a range of optics suited for UV range of laser light. The

beam was first oriented towards the region of interest in the test section using three 25.4

mm diameter mirrors. Then the beam diameter was increased and collimated using

spherical lenses. Two 25.4 mm diameter spherical lenses, a concave lens with f = -30 mm

and a convex lens with f = 50 mm were placed next to each other with an effective focal

length of f = -75 mm. The effective concave lens diverged the beam and increased its

diameter. A 50.8 mm diameter spherical convex lens with f = 300 mm was placed 205

mm ahead of them. This resulted in a collimated beam with an increased diameter of 40

mm. This beam was then converted to a sheet of light using cylindrical lenses. Two 50.8

x 50.8 mm cylindrical lenses were used, a convex lens with f = 300 mm and a concave

lens with f = -300 mm. The concave lens was placed 75 mm ahead of the convex lens,

and together they formed an effective convex lens, with the step placed 865 mm away at

the focal point. The effective cylindrical lens converged the beam in the horizontal

direction while leaving it unchanged in the vertical direction. The beam thus

progressively converged into an ellipse with diminishing minor axis and unchanged

major axis, and it formed a sheet 40 mm wide and less than 0.5 mm thick at the step. The

sheet thickness did not change appreciably in the recirculation region due to high

effective focal length of the lenses.









For image acquisition of acetone fluorescence in the visible spectrum, a Cooke

corporation intensified CCD camera and its associated software was used. The camera

has a resolution of 1280 x 1024 pixels, 12-bit dynamic range and shutter speed down to 3

ns. The software was used to set the camera parameters before taking images and also to

save the captured images on the computer. The images were captured with an exposure

time of 300 ms and a gain of 65%, with 100% representing the maximum gain

achievable. A 2 x 2 inning was performed on the image in horizontal and vertical

direction, that is, the average intensity of a square comprising 4 pixels was represented as

the value for one pixel that replaced the 4 pixels. The camera resolution was 12

pixels/mm.

Image Acquisition and Processing

For each image acquisition, a total of 20 images were captured to get a time-

averaged image over 2 s. Three set of images were required for analyzing each

experimental condition: the background image [Figure 2-5 (a)], the laser sheet profile

image [Figure 2-5 (b)], and the actual experiment image. To obtain the laser sheet profile,

the test section was closed and filled uniformly with acetone vapor, and then excited by

laser sheet. Since the acetone concentration was uniform, the acquired image captured the

spatial intensity variation in the laser sheet due to its Gaussian profile.

The nature of the experiment created challenges in getting a low background

image. The air arrived at the step base with M = 1.6 and had a low static temperature of T

= 200 K. However, it slowed down in the recirculation region to low subsonic Mach

numbers and the static temperature recovered close to the stagnation temperature of 300

K. Hence a temperature gradient existed in the shear layer separating the recirculation

region and the main airflow. This low air temperature caused the acetone vapor injected









in the test section to condense in the shear layer. The liquid acetone sprayed onto the

glass window and created a high background for the camera. This issue was overcome by

using a heat gun to heat the glass window up to 365 K just before starting airflow through

the test section. Even as the heat gun was in operation during the operation of the wind

tunnel, the glass window temperature dropped rapidly due to contact with cold airflow.

Hence fuel seeded with acetone was injected in the test section and images were acquired

within the first few minutes of starting the wind tunnel so that the glass window

temperature would not drop below 329 K, the boiling temperature for acetone. This

procedure reduced the background quite substantially. The background image set was

acquired just after the experiment was over and the airflow and heat gun were turned off.

The post processing of images was performed using a program written in Matlab.

The software is attached in Appendix B. An image is handled by Matlab as a 2D matrix

containing light intensity at each pixel, hence image processing is essentially an operation

with matrices. Each set of laser sheet profile, background and actual experiment images

was averaged to get a time-averaged image. Then the background image was subtracted

from the experiment image. The ultraviolet signal from Rayleigh scattering by acetone or

from Mie scattering by condensed acetone droplets was filtered by the camera lens since

it transmits only the visible spectrum of light. The laser sheet profile image was used to

correct the experiment image for spatial non-uniformity in laser intensity using the

following formula.


Icorrected (, n) = uncorrected (,) max[er( )] (2-1)
Laser (m, p)

where I: pixel intensity

m, n: arbitrary row and column location of a pixel in the image









p: a fixed column in the laser sheet image

With the exception of airflow shear layer and sonic fuel jet injected in the test

section, the temperature distribution in the recirculation region is rather uniform. Further,

the acetone LIF signal does not vary with temperature for low temperature range of 200-

300 K63, 64, 65. The pressure distribution in the recirculation region is rather uniform. The

acetone signal does not get quenched by oxygen62. Hence the LIF signal intensity

variation in the corrected image is independent of pressure, temperature and oxygen

concentration; it varies only with the concentration of acetone. With the assumption that

acetone distribution in the flow is the same as that for the fuel, the fuel mole fraction

distribution in the flow was determined quantitatively. The acetone LIF signal at the fuel

injection location, where fuel mole fraction is 100 %, was taken as the reference point for

other pixels in the image. The fuel mole fraction at a given pixel was then determined by

the ratio of LIF intensity at that pixel to the LIF intensity at the reference point.


Xfe(%)= *100 (2-2)
reference

For each experimental condition, image averaging was performed for 3

repeatability experiments. The fuel mole fraction distribution image, which does not

depend on the laser intensity variation, was used to determine the average PLIF image

and standard deviation for repeatability tests. The temporal variation in laser sheet profile

was quantified for the 60 min duration over which repeatability experiments were

performed. The image processing software is attached in Appendix B. Figure 2-6 shows

the laser sheet profile aty/H = 1.1 for three time intervals after the laser was started. Also

included is the profile when the laser was shut down and restarted. The figure shows only

a marginal change in laser profile over 85 min, the average standard deviation was 4%.









Restarting the laser did not affect the profile. The spikes in the profile reflect the quality

of the beam coming out of the laser.

Convolution filtering was applied on the average PLIF image to smoothen the

gradients due to noise. A 3 x 3 filter was applied to the image, that is, the signal intensity

at a given pixel was replaced by the average intensity in a 3 x 3 pixel square with the

given pixel at its center. The fuel mole fraction distribution was obtained from the filtered

image.

The binary diffusion coefficient of a gas in air is inversely proportional to the

square root of its molecular weight; it indicates that helium diffuses about 4 times faster

in air compared to acetone and argon diffuses at about the same rate as acetone. Hence

acetone will not be able to trace helium accurately while it will still trace argon

reasonably well. It suggests that the fuel mole fraction measurements determined from

acetone PLIF may be more accurate for argon as the injectant than helium. Condensation

of acetone vapor into droplets may also affect the PLIF signal intensity.






50




Table 2-1. Mass spectrometer calibration factors for various gases






















Mass sampling
ports and cooling
jacket


- Test section


Base
injection

- Upstream
injection

- Isolator



- Nozzle


------



Laser sheet
(PLIF)


i.--------------






Test section
Centerline








Ma, =1.6


MS-I
comp;


Inflow sampling
(Non-reacting flow)

Base injection
-- -- --






Wall sampling

Inflo% sampling
(Reacting flow)


L Upstream injection
.... M ,=l


Figure 2-1. Description of the test section showing fuel injection and mass sampling
locations: (a) image and (b) schematic diagram.














Shear Species Solenoid
layer valves
Recirculation -- o L- Manifold
region Cooling
jacket
Base fuel
injection "
I Water out
Supersonic
airflow Mass
spectrometer





Species composition
in recirculation region

(a)




Solenoid
Electronics valves


Inlet
Vacuum
chamber
Diffusion
Turbo PU
pump
pump




(b)



Figure 2-2. Mass sampling from the recirculation region behind the step for analysis by
the mass spectrometer: (a) schematic diagram (b) image.














4.0

3.5

3.0

2.5

S2.0

1.5

1.0

0.5

0.0


0 20 40 60


50

45

40

35

30

25

20

15

10

5
0


80 100 120 140
time (sec)

(a)


H2
-H20
- 02


0 20


40 60 80 100 120


time (sec)


(b)



Figure 2-3. MS measurements of species mole fraction distribution in the recirculation
region as the solenoid valves switch sequentially from one sampling port to
another under steady experimental conditions (a) non-reacting flow with
helium as the simulant fuel (b) reacting flow with hydrogen combustion.


- I /"\


F-1


RH --







54


Spherical


Computer


Figure 2-4. Schematic diagram of planar laser-induced fluorescence (PLIF) setup.































(a)
























(b)


Figure 2-5. Sample PLIF image for (a) background (b) laser sheet.



































0.5 1.0 1.5 20
x/H

Figure 2-6. Temporal variation of laser sheet profile aty/H = 1.1.














CHAPTER 3
RESULTS: NON-REACTING FLOW

The study included non-reacting cases with the air conditions maintained at M =

1.6, Toair = 300 K, Poair = 4.8 atm. Helium and argon were injected as fuel simulants at

two different pressures from the base of the step and from a location upstream of the step;

the configuration is described in Chapter 2. The molecular weight of helium is close to

hydrogen, while argon has molecular weight close to propane. The data obtained from

non-reacting flow tests is limited since combustion of fuel-air mixture does not take

place. However, it gives the fuel distribution in the flameholding region just before

ignition of fuel-air mixture. Based on this information from non-reacting flow, the effect

of various parameters, such as fuel injection pressure and location, on flameholding

characteristics in actual combustion can be estimated. The mass exchange rate across the

main airflow shear layer is higher for non-reacting flow compared to reacting flow24. For

non-reacting flow, this will bring more air into the recirculation region, hence the

flameholding region may be leaner in fuel compared to reacting flow.

Mass Spectrometry (MS)

Base Injection: Helium

Non-reacting flow experiments were performed with helium as the fuel simulant.

Each experiment was performed three times for repeatability. The average standard

deviation in XHe was 4 %. Mass sampling of the recirculation region species was done

along the wall in the x-direction with helium injected at the base of the step in Poair = 4.8

atm airflow. The wall pressure distribution for airflow without fuel injection is shown in









Figure 3-1. The supersonic airflow expands at the step base, as indicated by the sharp

drop in pressure at that location. The shear layer formed due to separation of the airflow

boundary layer at the step base is pushed towards the wall and it reattaches downstream

of the step. An oblique shock is formed at the reattachment point, causing a pressure rise

at that location. Based on the pressure rise in the plot, the shear layer reattachment point

is estimated to be 1.5-2.0 H. Oblique shocks result in a pressure rise towards the end of

test section to match the ambient atmospheric pressure. The pressure distribution in the

recirculation region remained unaffected due to fuel injection in the non-reacting flow

tests. It indicates that fuel injection did not cause a substantial increase in the mass flow

rate in the recirculation region.

Helium was injected at two pressures, a moderate stagnation pressure POHe = 5.4

atm and a high stagnation pressure POHe = 12.0 atm. The wall distribution of XHe in the

recirculation region for the two POHe is shown in Figure 3-2 (a). The standard deviation

bars are shown along with the average mole fractions for the repeated experiments. The

fuel injection location is indicated on the horizontal axis. It is observed that the fuel mole

fraction decreases in the x-direction and away from the fuel injection location, especially

for lower POHe. Increasing POHe substantially results in a corresponding increase in XHe in

the recirculation region. The XHe distribution shows more non-uniformity at higher POHe.

Mass sampling of the recirculation region species was done in the inflow z-

direction for the same airflow and fuel injection conditions as in wall sampling. The

inflow sampling was done at x/H = 2.0, y/H = 0.3. The inflow distribution of XHe in the

recirculation region is shown in Figure 3-2 (b) for the two POHe. At both pressures, the

inflow XHe are much higher than the wall measured XHe, specifically, up to 4-5 times.









However, XHe distribution away from the wall is rather uniform, indicating a well mixed

fuel-air mixture at the axial location probed.

Table 3-1 indicates the locally measured and the global XHe for wall and inflow

sampling. The global XHe, defined in the earlier chapters, is obtained from the total moles

of helium injected in the test section and the total moles of air flowing through the test

section. For both POHe, the locally measured XHe is up to 3 times more than the global

estimate for wall sampling and about 10 times more than the global estimate for inflow

sampling. It shows that even as the global XHe suggests a fuel-lean mixture, a fuel rich

mixture can exist in the recirculation region.

Helium has a molecular weight close to hydrogen, hence the XHe distribution

obtained from wall and inflow sampling is approximated to be similar to that of hydrogen

injected under identical test conditions. Using this assumption, if hydrogen had been

injected, the local equivalence ratios of hydrogen (PH2) could be determined from the

local mole fraction measurements. For POH2 = 5.4 atm, the local (nH2 is estimated in the

range of 0.04 0.4 for global (PH2 = 0.04. For POH2 = 12.0 atm, the local (PH2 is estimated

in the range of 0.1 0.7 for global (PH2 = 0.1. This confirms that a richer fuel composition

mixture exists in the recirculation region even as the global (H2 suggests a relatively fuel-

lean mixture. It is expected since only a fraction of the main airflow is entrained into the

recirculation region. Hence the global (H2 does not capture the actual non-premixed

conditions existing in the flameholding recirculation region.

Base Injection: Argon

Argon was injected at the base of the step for identical airflow and fuel injection

conditions as in the case of helium injection. The average standard deviation in XA was 3

%. Mass sampling of the recirculation region species was done along the wall in the x-









direction. The wall distribution of Xr in the recirculation region for the two PoAr is shown

in Figure 3-3 (a). The Xar distribution pattern is similar to that for helium injection,

especially for low injection pressure.

Mass sampling of the recirculation region species was done in the inflow z-

direction for the same airflow and fuel injection conditions as in wall sampling. The

inflow distribution of Xr in the recirculation region for the two PoAr is shown in Figure 3-

3 (b). As in wall sampling, the inflow XAM distribution has a pattern similar to that for

helium injection [Figure 3-2 (b)], especially for low injection pressure. The inflow Xa

measurements are up to 2-3 times higher than the wall measured Xr.

The locally measured and the global XAr for wall and inflow sampling are shown in

Table 3-2. The local to global Xfuel ratio can be interpreted as the fuel mole fraction in the

recirculation region for a unit mole fraction of fuel injected in the test section. A

comparison of data in Table 3-2 with Table 3-1 shows that the local to global Xfuel ratio is

higher for argon than helium, that is, higher concentration of argon is found in the

recirculation region than helium for a unit mole fraction of fuel injected in the test

section. This could be attributed to diffusion; the binary diffusion coefficient of a gas in

air is inversely proportional to the square root of its molecular weight, hence argon

diffuses about 3 times slower in air compared to helium.

Argon has a molecular weight close to propane, hence the Xar distribution obtained

from wall and inflow sampling is approximated to be same as that of propane injected

under identical test conditions. For POC3H8 = 5.4 atm, the local (PC3H8 is estimated in the

range of 0.5 2.5 for global (PC3H8 = 0.1. For POC3H8 = 12.0 atm, the local (PC3H8 is

estimated in the range of 1.0 4.3 for global (PC3H8 = 0.2. This shows that a fuel-rich









mixture exists in the recirculation region even as the global PC3H8 suggests a fuel-lean

mixture.

Upstream Injection: Helium

Helium was injected upstream of the step in identical airflow conditions as the base

injection experiments. The average standard deviation in XHe was 6 %. Helium was

injected at two pressures, POHe = 2.4 atm and 5.1 atm. The corresponding dynamic

pressure ratios [Pdynamic =PHeVHe2 / PairVair2] are 0.5 and 1.0 respectively. Mass sampling

of the recirculation region species was done along the wall in x-direction and in the

inflow z-direction. The wall distribution of XHe in the recirculation region for the two POHe

is shown in Figure 3-4 (a) and the inflow distribution is shown in Figure 3-4 (b). The

inflow sampling shows an almost proportional increase in XHe with POHe. Table 3-3

summarizes the local and global XHe for wall and inflow sampling. The plots and the table

show quite low levels of XHe in the recirculation region, hence indicating that upstream

injection of helium is not effective in supplying fuel to the recirculation region. In fact,

the local XHe is lesser than global XHe, indicating that the light gas penetrates through the

main airflow shear layer and only a small quantity reaches the recirculation region.

The XHe distribution obtained in the experiments is approximated to be the same as

that of hydrogen injected under identical test conditions. For POH2 = 2.4 atm, the local (pH2

is estimated in the range of 0.01 0.02 for global (PH2 = 0.02. For POH2 = 5.1 atm, the local

(PH2 is estimated in the range of 0.01 0.03 for global (PH2 = 0.05. The fuel-lean

conditions for upstream injection mode suggest difficulty in flameholding in a

combustion experiment with hydrogen as the fuel.









Upstream Injection: Argon

Argon was injected upstream of the step for identical airflow and fuel injection

conditions as in case of helium injection. Hence, argon was injected at the same dynamic

pressure ratios as helium. The average standard deviation in XM was 3 %. Mass sampling

of the recirculation region species was done along the wall in x-direction and in the

inflow z-direction. The wall distribution of XA in the recirculation region for the two PoAr

is shown in Figure 3-5 (a) and the inflow distribution is shown in Figure 3-5 (b). Table 3-

4 summarizes the local and global XAr for wall and inflow sampling. Unlike the case of

helium, the heavier gas argon reaches the recirculation region in quantities larger than the

global XA. It can be seen from the plots and from the local to global XAr ratios in the table

that an increase in the upstream PoAr does not result in a corresponding increase in the

amount of fuel reaching the recirculation region. If fuel is injected upstream at a low

dynamic pressure ratio, it seeps into the boundary layer of the incoming airflow which

carries it into the recirculation region. If the upstream fuel injection dynamic pressure

ratio is increased, part of the fuel penetrates through the airflow boundary layer and

escapes into the core airflow; hence less fuel reaches the recirculation region.

The XAM distribution obtained in the experiments is approximated to be the same as

that of propane injected under identical test conditions. For POC3H8 = 2.4 atm, the local

(PC3H8 is estimated in the range of 0.2 0.3 for global (PC3H8 = 0.06. For POC3H8 = 5.1 atm,

the local (PC3H8 is estimated in the range of 0.2 0.3 for global (PC3H8 = 0.1. The fuel-lean

conditions in the recirculation region for upstream injection mode, along with much

lower flameholding limits for hydrocarbons as compared to hydrogen suggest difficulty

in flameholding in a combustion experiment with propane as the fuel.









Planar Laser Induced Fluorescence (PLIF)

Base Injection: Helium

PLIF imaging of the recirculation region fuel distribution in the z/W = 0.9 plane

was performed for identical airflow and fuel injection conditions as the mass sampling

experiments. Each experiment was performed 3 times for repeatability. The average

standard deviation in XHe, after deducting the XHe deviation due to the temporal variation

in laser sheet profile discussed in Chapter 2, was 8 %. The results for helium injection are

shown in Figures 3-6 and 3-7 for POHe = 5.4 atm and 12.0 atm respectively. The PLIF

image for POHe = 5.4 atm is shown in Figure 3-6 (a) and the XHe distribution is shown in

Figure 3-6 (b). The expansion of airflow at the step pushes the shear layer towards the

test section wall. The fuel injection holes are inclined relative to the step base. This is

clearly visible in the figures as the fuel jet impinges on the step base and forms a plume

above it. This fuel injection configuration helps the fuel remain and mix within the

recirculation region. The PLIF image for POHe = 12.0 atm is shown in Figure 3-7 (a) and

the XHe distribution is shown in Figure 3-7 (b). In agreement with the mass spectrometer

measurements, the fuel remains in the recirculation region even as the injection pressure

is increased. The XHe distribution shows more non uniformity for higher injection

pressure.

The global XHe for POHe = 5.4 atm and 12.0 atm are 1.1 % and 2.5 % respectively.

As in mass spectrometer measurements, the local XHe in the recirculation region and shear

layer [Figure 3-6 (b) and 3-7 (b)] are an order of magnitude higher than the

corresponding global XHe values. The XHe distribution obtained in the experiments is

approximated to be the same as that of hydrogen injected under identical test conditions.

For POH2 = 5.4 atm, the global PH2 = 0.04 and for POH2 = 12.0 atm, the global (PH2 = 0.1.









The global (H2 values suggest a fuel lean mixture for both injection pressures. With the

observation that (H2 = 1 corresponds to 30 % XH2 in a hydrogen-air mixture, it is seen in

Figure 3-6 (b) that for POH2 = 5.4 atm, part of the recirculation region has a fuel-rich

mixture with (H2 > 1. Figure 3-7 (b) shows that for POH2 = 12.0 atm, the entire

recirculation region, excluding the main airflow shear layer, has (H2 > 1.

Base Injection: Argon

Argon was injected at the base of the step for identical airflow and fuel injection

conditions as in the case of helium injection. The average standard deviation in XA, after

deducting the XA deviation due to the temporal variation in laser sheet profile, was 8 %.

The results for argon injection are shown in Figures 3-8 and 3-9 for PoA = 5.4 atm and

12.0 atm respectively. The PLIF image for PoAr = 5.4 atm is shown in Figure 3-8 (a) and

the XA distribution is shown in Figure 3-8 (b). The PLIF image for Po0 = 12.0 atm is

shown in Figure 3-9 (a) and the XA distribution is shown in Figure 3-9 (b). As in the case

of helium injection, the fuel remains within the recirculation region for both injection

pressures. By comparing fuel mole fraction distribution in Figure 3-8 (b) with Figure 3-6

(b) for low injection pressure and Figure 3-9 (b) with Figure 3-7 (b) for high injection

pressure, it is observed that argon distribution in the recirculation region is similar to

helium, especially for low injection pressure. This is in agreement with the mass

spectrometer measurements.

The global XA for PoAr = 5.4 atm and 12.0 atm are 0.4 % and 0.8 % respectively.

As in helium injection, a much richer fuel composition exists in the recirculation region

[Figure 3-8 (b) and 3-9 (b)] than suggested by global XA values. The XA distribution

obtained in the experiments is approximated to be the same as that of propane injected

under identical test conditions. For POC3H8 = 5.4 atm, the global (PC3H8 = 0.1 and for POC3H8









= 12.0 atm, the global PC3H8 = 0.2. The global PC3H8 values suggest a fuel lean mixture for

both injection pressures. With the observation that (PC3H8 = 1 corresponds to 4 % Xc3H8 in

a propane-air mixture, it is seen in Figures 3-8 (b) and 3-9 (b) that for both injection

pressures, the entire recirculation region and the main airflow shear layer is fuel-rich in

composition, with PC3H8 > 2.

Upstream Injection

PLIF imaging of the recirculation region fuel distribution was performed for

identical airflow and fuel injection conditions as the mass sampling experiments. For all

test conditions, the PLIF signal was barely noticeable above the background. It is in

agreement with mass spectrometer measurements and shows that hardly any fuel reaches

the recirculation region for upstream injection.

Comparison between MS and PLIF data

The fuel mole fraction measurements at x/H = 2.0, y/H = 0.3, z/W = 0.9 obtained

from PLIF and MS are compared in Figure 3-10 for base fuel injection of helium [Figure

3-10 (a)] and argon [Figure 3-10 (b)]. For PLIF, the fuel mole fractions obtained from

unfiltered image and averaged over the mass sampling tube cross-section are reported.

Filtering increased the fuel mole fraction by 10 % for helium and 2 % for argon injection,

hence unfiltered data was used for comparison.

The PLIF and MS data do not overlap with each other for all test conditions. One of

the reasons for the difference between PLIF and MS measurements is the intrusive nature

of mass sampling for MS measurements. The outer diameter of the 3 inflow sampling

tubes was 0.9 mm; the small diameter tubes were selected to minimize disturbances to the

flow field. However, changes in the local flow field due to the presence of sampling tubes

cannot be eliminated completely. A more significant disturbance in the local flow field









was caused by suction applied at the sampling port tip to extract gas mixture from the

recirculation region into the mass spectrometer; the pressure gradient was high, with

about 0.5 atm in the flow field behind the step and near vacuum at the entrance of the

mass spectrometer. Hence, a difference between MS and PLIF data is expected due to

different local flow fields in the two measurements.

As discussed in Chapter 2, the limitation of PLIF measurements is that acetone is

expected to trace argon better than helium due to the difference in diffusion rates.

Acetone diffuses about 4 times slower than helium and about the same rate as argon.

Hence acetone PLIF measurements will over-estimate helium mole fraction in the flow.

This is observed in the MS-PLIF data comparison; the difference between acetone PLIF

fuel mole fraction measurements and corresponding MS measurements is higher for

helium compared to argon.

The difference could also be due to a slight misalignment of the measurement point

between the two techniques due to shaking of the test section as air flowed through it.

The point at which PLIF and MS data are compared lies in the airflow shear layer. In this

region, sharp gradients exist in the x-y plane, even though the flow is rather uniform in

the z direction as seen in inflow MS measurements. The lateral test section movement

observed by the camera during PLIF tests was 3 pixels (Ay / H = 0.02). For POHe = 5.4

atm and 12.0 atm test conditions, a lateral shift in measurement location from the

reference point at y/H = 0.3 by Ay / H = 0.02 results in a 3 % change in fuel mole

fraction. Hence shaking of the test section is not a significant reason for the difference

between MS and PLIF data.









Table 3-1. Base fuel injection: global and local XHe.


PO He (atm) local XHe() global XHe %) local / global XHe
Wall sampling
5.4 1.1-3.8 1.1 0.9-3.4
12.0 2.6-8.3 2.5 1.0-3.4
Inflow sampling
5.4 2.2-11.8 1.1 1.9-10.4
12.0 5.7-21.5 2.5 2.3-8.7

Table 3-2. Base fuel injection: global and local Xr.


Po (atm) local XA (%) global XA (%) local / global XA
Wall sampling
5.4 2.0-5.0 0.4 5.5-13.9
12.0 3.8-9.9 0.8 4.9-12.7
Inflow sampling
5.4 2.9-9.9 0.4 7.9-27.4
12.0 5.2-17.1 0.8 6.6-21.9

Table 3-3. Upstream fuel injection: global and local XHe.


Po He (atm) local XHe (% global XHe (%) local / global XHe
Wall sampling
2.4 0.2-0.3 0.7 0.3-0.5
5.1 0.3-0.4 1.5 0.2-0.3
Inflow sampling
2.4 0.3-0.6 0.7 0.4-0.9
5.1 0.4-1.0 1.5 0.2-0.7

Table 3-4. Upstream fuel injection: global and local XM.


Po (atm) local XA (%) global XA (%) local / global XA
Wall sampling
2.4 0.7-0.9 0.2 3.3-4.0
5.1 0.8-1.0 0.5 1.7-2.0
Inflow sampling
2.4 0.7-1.0 0.2 3.3-4.5
5.1 0.9-1.3 0.5 1.9-2.8



















S
S


S
S
~ilffi~


-20 -15 -10 -5


0 5 10 15 20 25

x/H


Figure 3-1. Wall pressure distribution for non-reacting flow. Pair = 4.8 atm, Mair = 1.6.
The axial origin is placed at the step.





















- POHe 5.4 atm
- PoH = 12.0 atm


- 4 -


0.0 .5 1.0 1.5 2.0 2.5 3.0 3.5 4.0
x/H


y


(a)
(a)


- POHe 5.4 atm
/ POHe = 12.0 atm


0.0 0.2 0.4 0.6
z/W


0.8 1.0


z


x
(b)




Figure 3-2. Base fuel injection: MS measurement of helium mole fraction distribution in
the recirculation region for (a) wall sampling (b) inflow sampling. Poair = 4.8
atm, Mair = 1.6.


25

20

15

10

5
n


u















20

-15

10

5

n


PAr = 5.4 atm
- PAr 12.0 atm


0.0 A.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0
x/H

~ --. -(a )---- ty

x

(a)


20

515

10

5

0


ooL
/ I


0.0 0.2


0.4 0.6
zlW


POAr = 5.4 atm
POAr = 12.0 atm


0.8 1.0


z

x

(b)




Figure 3-3. Base fuel injection: MS measurement of argon mole fraction distribution in
the recirculation region for (a) wall sampling (b) inflow sampling. Pair = 4.8
atm, Mair = 1.6.


0













1.2
1.0
0.8
POHe 2.4 atm
S0.6 POH, 5.1 atm
X 0.4 .-.
0.2
0.0
0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0
x/H

ty

x
(a)


1.2
1.0
-0.8
S0.6
X 0.4
0.2
0.0


POHe =2.4 atm
- POHe 5.1 atm


0.0 0.2 0.4 0.6 0.8 1.0
z/W


x


Figure 3-4. Upstream fuel injection: MS
distribution in the recirculation
sampling. Pair = 4.8 atm, Mair = 1


measurement of helium mole fraction
region for (a) wall sampling (b) inflow























Po = 2.4 atm
--- Po =5.1 atm
I -- --*-. --- i


0.0 i
0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0
x/H




x

(a)


a--
- -


Po = 2.4 atm
-. Pom =5.1atm


0.0 0.2 0.4 0.6 0.8 1.0
z/W
z


x

(b)


Figure 3-5. Upstream fuel injection: MS measurement of argon mole fraction
distribution in the recirculation region for (a) wall sampling (b) inflow
sampling. Pair = 4.8 atm, Mair = 1.6.


1.5

1.0


2.0

1.5

1.0

0.5

0.0





















Fuel


0.2H T


I Airflow


1.5



1.0



0.5


(b)
Figure 3-6. PLIF measurement for base injection of helium (a) image (b) XHe
distribution (%). POHe = 5.4 atm, Poair = 4.8 atm, Mair = 1.6.
























Fuel



0.2 H


I Airflow


1.5




S1.0




0.5


1.0 1.25


Figure 3-7. PLIF measurement for base injection of helium (a) image (b) XHe
distribution (%). POHe = 12.0 atm, Poair = 4.8 atm, Mair = 1.6.





























Fuel




0.2 HT


Airflow


2.0




1.5




I
- 1.0




0.5


1.0 1.25


Figure 3-8. PLIF measurement for base injection of argon (a) image (b) XA distribution
(%). PAr = 5.4 atm, Poair = 4.8 atm, Mair = 1.6.






















Fuel



0.2 H


T Airflow


(b)
Figure 3-9. PLIF measurement for base injection of argon (a) image (b) XA distribution
(%). PoAr = 12.0 atm, Poair = 4.8 atm, Mair = 1.6.























5 '*

I ------------



0.0 0.2 0.4 0.6 0.8 1.0
z/W


POHe 5.4 atm(MS)

POHe 12.0 atm (MS)

POHe 5.4 atm (PLIF)

POHe 12.0 atm (PLIF)


20

-15

10

5

0


0 e



*^ >^ i


Po, 5.4 atm (MS)

PoAe 12.0 atm (MS)

PoAe 5.4 atm (PLIF)

Poe = 12.0 atm (PLIF)


0.0 0.2 0.4 0.6 0.8 1.0
z/W

(b)




Figure 3-10. Comparison between MS and PLIF data for base fuel injection of (a) helium
(b) argon.














CHAPTER 4
RESULTS: REACTING FLOW

The study included reacting cases with the air conditions maintained at M = 1.6,

Toair = 300 K, Poair = 4.5 atm. Hydrogen was injected as fuel at two different pressures

from the base of the step and from a location upstream of the step; the configuration is

described in Chapter 2. The results are described below.

Base Injection: Hydrogen

Combustion experiments were performed with hydrogen injected at the base of the

step. Mass sampling of the recirculation region species was done along the wall in the x-

direction. Hydrogen was injected at two stagnation pressures: POH2 = 4.5 atm and POH2 =

8.2 atm, the air stagnation pressure was Poair = 4.5 atm. These corresponded to global (pH2

of 0.04 and 0.08 respectively. The wall pressure distribution for the two POH2 is shown in

Figure 4-1. For POH2 = 4.5 atm, the expansion of airflow approaching the step is reduced

due to heat released from combustion and pressure rise at the step base. Hence the

recirculation region length is more than the corresponding non-reacting case. For POH2 =

8.2 atm, the rather uniform pressure distribution due to more heat release implies an even

longer recirculation region and probably no reattachment point for the shear layer. Figure

4-2 (a) shows the wall distribution of local PH2 in the recirculation region for POH2 = 4.5

atm. Figure 4-3 (a) shows the corresponding plot for POH2 = 8.2 atm. The local (pH2 was

deduced from the mole fractions of hydrogen and water in the product mixture. The fuel

injection location is indicated on the horizontal axis in the plots. A highly non-uniform

(PH2 distribution is observed in the recirculation region with a maximum around 2.2H. The









local (PH2 goes up to 0.7 at POH2 = 4.5 atm and up to 1.3 at POH2 = 8.2 atm, showing a

proportional increase due to an increase in POH2.

The wall distribution of products from hydrogen combustion for the two POH2 is

shown in Figures 4-2 (b) and 4-3 (b). XN2 has not been included in the plots. Concurrent

with the earlier observation of a fuel-rich mixture existing in the recirculation region, the

combustion product composition shows a significant proportion of unburned hydrogen.

XH20 increases as we move downstream of the step. However a significant proportion of

unburned oxygen and only a small proportion of water are a reflection of the limitation of

sampling at the wall where the combustion radicals get quenched and hence the

composition can be quite different from elsewhere in the flow.

Inflow mass sampling of the recirculation region species was done for the same

airflow and fuel injection conditions as in wall sampling. Figures 4-4 and 4-5 show the

distribution of local (PH2 and product mole fractions in the recirculation region for POH2 =

4.5 atm and POH2 = 8.2 atm respectively. The experimental conditions were not identical

over the repeated tests, both Poair and POH2 varied by 0.3 atm. Hence the high standard

deviations observed in the plots could be attributed to the change in recirculation region

mixture composition due to changes in airflow and fuel injection conditions. The local

(PH2 distributions in Figures 4-4 (a) and 4-5 (a) show a decreasing amount of fuel as we go

downstream of the fuel injection location. The local pH2 does not increase proportionally

with the increase in POH2. As POH2 increased, the local (PH2 increased unevenly in the

recirculation region; the region close to the injection location experienced a lower

increase in PH2 than the region farther away in the x-direction. Thus, increasing POH2 led

to a reduction of local gradients in the recirculation region. The product mole fraction









distributions in Figures 4-4 (b) and 4-5 (b) show a fuel rich mixture in the recirculation

region with plenty of unburned hydrogen and almost no oxygen. The proportion of

unburned hydrogen drops rapidly as we go downstream in the recirculation region. For

low POH2, once the hydrogen was completely consumed, the oxygen mole fraction

increased. For high POH2, hydrogen was still present at the last sampling port and oxygen

was virtually nonexistent. The temperature drop of the sampled mixture while passing

through the cooling jacket resulted in condensation of water vapor and much of it could

not reach the mass spectrometer. Hence the corrected XH20 was deduced from the oxygen

deficit in the product mixture. Both corrected and uncorrected XH20 are plotted in the

figures. Unlike the wall sampling experiments, significant amount of water was produced

and XH20 increases as we go downstream in the recirculation region.

The local and global (PH2 for the two POH2 obtained from wall and inflow samplings

are compared in Table 4-1. As in the non-reacting flow test results discussed earlier, more

fuel is observed away from the wall. For both POH2, the local PH2 is an order of magnitude

higher than the suggested global value.

Upstream Injection: Hydrogen

Hydrogen was injected upstream of the step for identical airflow conditions as base

injection. Hydrogen was injected at two stagnation pressures: POH2 = 2.5 atm and POH2 =

8.2 atm. The corresponding dynamic pressure ratios [Pdynamic = PH2VH22 / airVair2] are 0.5

and 1.6 respectively. However, a flame could not be established for both POH2, which is in

line with the predictions of non-reacting flow experiments with helium injected upstream

of the step. Mass sampling of the recirculation region species along the wall in x-

direction showed a fuel-lean mixture with no water formed as a byproduct of combustion.






81


A change in the injection configuration, e.g., number of orifices, angled injection, etc.,

may lead to possibly holding the flame.






82


Table 4-1. Base fuel injection: global and local OH2


PO H2 (atm) local (DH2 global (DH2 local/global (DH2
Wall sampling
4.5 0.1-0.7 0.04 1.8-18.3
8.2 0.2-1.3 0.08 2.6-16.4
Inflow sampling
4.5 0.8-2.7 0.04 21.0-66.5
8.2 1.5-2.8 0.08 18.5-35.0















A A A
S S
U_ *Sg@


* PH2= 4.5 atm
SPOH2 = 8.2 atm


-20-15-10 -5 0 5 10 15 20 25
x/H


Wall pressure distribution for hydrogen combustion tests. Pair = 4.5 atm,
Mair = 1.6.


1.4

1.2
1.0

S0.8
a-
0.6

0.4
a-
0.2
0.0


Figure 4-1.

















3.
2.
2.
M 1.
9-
1.
0.
0.







80
70
60
S50
40
SAn


0.0 0.5 1.0 1.5


2.0 2.5 3.0 3.5 4.0
x/H


(b)


Y
i----


Figure 4-2. Base fuel injection: Wall sampling results for (a) hydrogen equivalence
ratio distribution in the recirculation region (b) combustion species mole
fraction distribution. POH2 = 4.5 atm, Poair = 4.5 atm, Mair = 1.6.


0
5
0
5
0
5
0
.0 t ----------------
0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0
x/H

(a)











S---- --------


-. H2
- H20
- 02

















3.0
2.5
2.0
1.5
1.0
0.5
0.0


0.0 O.5 1.0 1.5


2.5 3.0 3.5 4.0


-, H2
- H20
- 02


/ ,

A c


0.0 0.5 1.0 1.5


2.0 2.5 3.0 3.5 4.0
x/H

(b)



--- y


Figure 4-3. Base fuel injection: Wall sampling results for (a) hydrogen equivalence
ratio distribution in the recirculation region (b) combustion species mole
fraction distribution. POH2 = 8.2 atm, Poair = 4.5 atm, Mair = 1.6.


-50
40
330
20
10
0




















3.0

2.5

2.0
C4
1.5

1.0

0.5


-K


U.L


0.0 .5 1.0 1.5 2.0 2.5 3.0 3.5 4.0
x/H


80
70
60
50
.-40
30
20
10
0
0


T



~--- N.---
r I


.0


- H2
- H20
- 02
- H20 uncorrected


p.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0
x/H


(b)





x


Figure 4-4. Base fuel injection: Inflow sampling results for (a) hydrogen equivalence
ratio distribution in the recirculation region (b) combustion species mole
fraction distribution. POH2 = 4.5 atm, Poair = 4.5 atm, Mair = 1.6.

















3.0
2.5
2.0
S 1.5
1.0
0.5


U.U
0.C


0 .5 1.0 1.5 2.0 2.5 3.0 3.5 4.0
x/H

(a)





\


S.. -..

.- ... --- *
T i i i i


80
70
60
50
40
30
20
10
0


, H2
. H20
2
- H20 uncorrected


0.0 0.5 10 1.5 2.0 2.5 3.0 3.5 4.0
x/H

(b)



z
x
_rrrr- Lr_

X


Figure 4-5. Base fuel injection: Inflow sampling results for (a) hydrogen equivalence
ratio distribution in the recirculation region (b) combustion species mole
fraction distribution. POH2 = 8.2 atm, Poair = 4.5 atm, Mair = 1.6.














CHAPTER 5
CONCLUSIONS

Mass spectrometry (MS) and planar laser induced fluorescence (PLIF) were used to

determine the species concentration distribution in the flameholding recirculation region

and free shear layer formed behind a rectangular step in supersonic flow. Non-reacting

and combustion tests were conducted and fuel related parameters such as the injection

location, injection pressure and fuel type were varied. The conclusions are summarized

below.

Fuel injection location:

Fuel injection at the base of the step was effective in supplying fuel directly into

the flameholding region. Stable flames were achieved in combustion tests.

Fuel injection upstream of the step was not effective in supplying sufficient

amount of fuel to the recirculation region, hence a flame could not be sustained in

combustion tests. An injection configuration change, such as increasing the

number of injection holes, could improve flameholding for upstream injection.

Fuel concentration in the recirculation region:

Base injection:

The local fuel concentration in the recirculation region was an order of

magnitude higher than the suggested global fuel mole fraction since only a small

part of the main airflow entered the recirculation region. For combustion tests

with hydrogen, the recirculation region was predominantly fuel-rich in

composition even for quite low global (pH2.









The above observation implies that for the same global equivalence ratio, the

flameholding region is richer in fuel composition for non-premixed case

compared to premixed case. As the global equivalence ratio is increased, the non-

premixed case will flood the recirculation region with fuel and hence have worse

fuel-rich flame stability limit than the premixed case. On the other hand, as the

global equivalence ratio is reduced, the non-premixed case will still have

sufficient fuel in the recirculation region and hence have a better fuel-lean

stability limit than the premixed case.

Non-reacting flow experiments indicated a leaner composition of the

recirculation region than the combustion tests. It could be due to higher air

entrainment rate into the recirculation region through the main airflow shear layer

for non-reacting flow tests than combustion tests.

For non-reacting and reacting flow tests, higher fuel concentration was

measured inflow in the recirculation region than at the test section wall.

Fuel injection pressure:

Base injection:

For non-reacting flow tests, fuel remained in the recirculation region and shear

layer for both injection pressures. The fuel injection holes are inclined relative to

the step base; such a configuration helps the fuel remain and mix within the

recirculation region. Fuel distribution in the recirculation region was more non-

uniform for higher fuel injection pressure.

For combustion tests, increasing the fuel injection pressure resulted in reducing

the fuel distribution gradient inflow in the recirculation region.









Upstream injection:

The lighter gas, i.e. helium, penetrated through the airflow boundary layer along

the wall for both injection pressures; the fuel quantity detected in the recirculation

region was less than the suggested global value.

At higher injection pressure for the heavier gas, i.e. argon, the jet penetrated

through the airflow boundary layer and less of it was carried into the recirculation

region. However, for both argon injection pressures, more fuel was detected in the

recirculation region than the indicated global value.

Fuel type:

Base injection:

The fuel distribution pattern in the recirculation region was similar for helium

and argon. However, argon concentration in the recirculation region was higher

than helium for a unit mole fraction of fuel injected in the test section. It could be

due to slower diffusion rate of argon in air than that of helium in air.

The above observation implies that as the fuel injection pressure is increased,

the heavier fuel will have higher fuel mole fraction in the recirculation region, and

hence have a worse fuel-rich flame stability limit, than the lighter fuel. For the

same reason, as the injection pressure is decreased, the heavier fuel will still have

enough fuel mole fraction in the recirculation region, and hence have a better fuel-

lean flame stability limit, compared to the lighter fuel. This analysis is in addition

to the fact that a heavier fuel such as propane has a much smaller flame stability

curve than a lighter fuel such as hydrogen.