As the flight speeds of aircraft are increased above Mach 5, efficient methods of propulsion are needed. Scramjets may be a solution to this problem. Supersonic combustion is one of the main challenges involved in the operation of a Scramjet engine. In general, both an igniter and a flameholder are needed to achieve and maintain supersonic combustion. The current work examines a plasma torch-strut combination as an igniter-flameholder. The plasma torch-strut combination was tested in the Virginia Tech unheated supersonic wind tunnel at Mach 2.4. Pressure and temperature sampling, filtered photography, and spectroscopic measurements were used to compare different test cases. These results provide both qualitative and quantitative results on how the combination responds to changes in the mass flow rate of fuel and the power to the plasma torch. The key conclusions of the work were the following: 1. Tests showed that an exothermic reaction takes place. 2. The amount of heat release increases with an increase in the mass flow rate of fuel. 3. The plasma torch-fuel injector interaction caused the heat release to be well above the tunnel floor and sometimes off the strut centerline 4. One change in the fuel injector pattern caused more temperature rise near the floor of the tunnel. 5. The flow penetration height of the plasma torch alone was reduced by the fuel-plasma torch interaction. 6. Moving the strut upstream reduced the measured temperature rise at a fixed downstream location, but increased the penetration height of the plasma torch. 7. The computed heat release was found to be small compared to the potential heat release from all the fuel burning. 8. The amount of temperature rise caused by the fuel is not greatly affected by the power to the plasma torch. / Master of Science

Scramjet engines are the focus of considerable interest for propulsion in the hypersonic flow regime. One of the serious technical challenges for developing scramjets is reducing the skin friction drag on the engine. The combustion chamber, in particular, is a major contributor to the skin friction drag because of the high density of the flow through that region. This investigation focuses on reducing the combustion chamber skin friction drag by minimising the surface area and size of the combustion chamber and by employing a novel approach to accomplishing combustion. The first design criterion is addressed by using a single internal-combustor scramjet configuration, as opposed to multiple external combustors, and by injecting the fuel on the intake to reduce the mixing length required in the combustor. The second design criterion refers to the use of a new technique called radical farming. This uses the highly two-dimensional nature of the flow through the engine, which is created by deliberately ingesting the leading edge shocks, to achieve combustion at lower mean static pressures and temperatures than generally expected. A simplified approximate theoretical analysis of the radical farming concept is presented. Experiments were conducted in the T4 free-piston shock tunnel on a scramjet model with a single rectangular constant cross-sectional area combustion chamber. Pressure measurements were taken along the centreline of the intake, combustion chamber and thrust surface and across the model width at three locations. Gaseous hydrogen fuel was injected halfway along the intake at a range of equivalence ratios between zero and one. The combustion chamber height was varied from 20mm to 32mm, which varied the contraction ratio of the engine from 4.1 to 2.9. The experiments were conducted at a stagnation enthalpy of either 3MJ/kg or 4MJ/kg. The nominal 3MJ/kg condition corresponds to Mach 7.9 flight at an altitude of 24km. The majority of the 4MJ/kg experiments were conducted at a nominal condition corresponding to Mach 9.1 flight at an altitude of 32km. A small number of 4MJ/kg experiments were conducted at simulated flight altitudes of between 30 and 38km; the flight Mach number for these experiments was approximately 9.0. Thrust was calculated by integrating the centreline pressure distribution over the area of the thrust surface, assuming that the pressure at any axial location was constant across the engine width. These experimental thrust values were compared with theoretical estimates obtained using a one-dimensional analysis and a quasi-two-dimensional analysis. The comparison provided an indication of the level of completion of combustion in the experiments. The difference in thrust produced as a result of combusting fuel was examined by plotting the incremental specific impulse against equivalence ratio. Experimental and theoretical results agreed best at the higher equivalence ratios. Turbulent boundary layer separation correlations were used to provide reasonable estimates for the equivalence ratio at which the flow choked. The drag on the internal flowpath of the scramjet engine was estimated using the quasi-two-dimensional analysis. This drag estimate was combined with the experimental thrust measurements to provide estimates of the net specific impulse. Positive net specific impulse estimates were obtained above a certain minimum equivalence ratio, which depended on the contraction ratio and the test condition. The engine performance was observed to be highly dependent on the two-dimensional shock structure within the engine. Thrust and specific impulse were observed to decrease with increasing simulated flight altitude, as expected. Positive net specific impulse estimates were obtained at equivalence ratios of approximately one for simulated flight altitudes below 35km. Assuming complete combustion and that an equivalence ratio of one can be reached, the configuration considered in the present study can theoretically reach a net specific impulse of approximately 1000s at the 3MJ/kg condition and 500s at the 4MJ/kg condition. These numbers provide a promising testimonial for the use of this configuration, with modifications, as a more efficient alternative to rocket engines.

A direct connect, supersonic solid fuel combustor with a cavity is explored in the context of understanding characteristics related to ignition, regression rate, combustion, and flow fields for application in advancing solid fuel scramjet research. 3D printed, polymethylmethacrylate fuel grains are loaded into both fully enclosed and optically accessible combustors. The ignition characteristics are investigated by systematically varying the internal geometry of the fuel grain to develop a flammability map with respect to non-dimensional geometric parameters. Results reveal that a longer and larger flameholding cavity creates favorable conditions for ignition and sustained combustion. The inlet temperature is also systematically varied to extend the available literature on the supersonic combustion of solid fuels to lower temperature operating conditions and show that a higher inlet temperature is conducive to sustained combustion and higher regression rates. The regression rates of the fuel grains are measured to determine a concentration of regression in the flameholding cavity along the angle of the downstream side of the cavity. Ignition and sustained combustion rely heavily on the fuel in the flameholding cavity. A decreasing regression rate is observed as the fuel regresses by measuring the regression rate at discrete time intervals during a firing of the optical combustor. The optical combustor is also subject to various high-frequency imaging techniques. Shadowgraph imaging shows the changes in density of the flow field and finds a normal shock in the constant area section. CH* chemiluminescence imaging provides novel observations of the concentrated areas of combustion along the fuel grain wall by highlighting the heat release from combustion. A high intensity of CH* radicals is in the upstream section of the flameholding cavity. When considered in the context of the concentration of regression, this indicates that the recirculation zone pulls fuel from the downstream section of the cavity, combusts it in the upstream section of the flameholding cavity, then expels the higher enthalpy gas into the core flow. Additionally, observing the flow provides insight into the flow dynamics of opposing cavities in a supersonic flow field. The symmetry of the flow field is found to be reliant on the stability of the flameholding cavity length to depth ratio. / Master of Science / A solid fuel scramjet has the potential to be the simplest and most cost effective method of achieving hypersonic flight. A liquid fuel scramjet has been demonstrated in free flight, but liquid fuels present many issues involving safety and storage that can be eliminated by introducing solid fuels. Supersonic combustion, or burning fuel in an air flow moving faster than the speed of sound, is a complicated subject due to the irregularity of flow fields and the requirement of combustion to occur at a high rate. The research within this thesis presents many novel technologies that have never been presented in published literature in the context of the supersonic combustion of solid fuels. By conducting ground testing of a solid fuel scramjet, characteristics of the combustion can be studied to expand the available literature in the field to new fuel geometries and inlet conditions. The ignition and sustained combustion of a solid fuel scramjet is extremely reliant on the initial geometry of the fuel and the initial temperature of the flow. This research advances the field of supersonic combustion of solid fuels by developing an optically accessible combustor using quartz windows. These characteristics of supersonic combustion are investigated using highspeed video recording. The results of these techniques provide insight into favorable fuel geometries and inlet conditions. Additionally, patterns observed in the flow field explain concentrations of combustion and fuel consumption.

A flush-wall injector model and a strut injector model representative of state of the art scramjet engine combustion chambers were experimentally studied in a cold-flow (non-combusting) environment to determine their fuel-air mixing behavior under different operating conditions. The experiments were run at nominal freestream Mach numbers of 2 and 4, which simulates combustor conditions for nominal flight Mach numbers of 5 and 10. The flush-wall injector model consists of sixteen inclined, round, sonic injectors distributed around the wall of a circular duct. The strut injector model has sixteen inclined, round, sonic injectors distributed across four struts within a circular duct. The struts are slender, inclined at a low angle to minimize drag, and have two injectors on each side. The experiments investigated the effects of injectant molecular weight, freestream Mach number, and jet-to-freestream momentum flux ratio on the fuel-air mixing process. Helium, methane, and air injectants were studied to vary the injectant molecular weight over the range of 4-29. All of these experiments were performed to support the needs of an integrated experimental and computational research program, which has the goal of upgrading the turbulence models that are used for Computational Fluid Dynamics predictions of the flow inside a scramjet combustor. The primary goals of this study were to use injector models that represent state of the art scramjet engine combustion chambers to provide validation data to support the development of turbulence model upgrades and to add to the sparse database of mixing results in such configurations. The main experimental results showed that higher molecular weight injectants had approximately the same amount of penetration in the far field as lower molecular weight injectants at the same jet-to-freestream momentum flux ratio. Higher molecular weight injectants also demonstrated a mixing rate that was the same as or slower than lower molecular weight injectants depending on the flow conditions. A comparison of the experimental results for the two different injector models revealed that the flush-wall injector mixed significantly faster than the strut injector in all of the experimental cases. / Ph. D.

No description available.

An integrated aerodynamic-ramp-injector/plasma-torch-igniter of original design was tested in a Mâ = 2, unvitiated, heated flow facility arranged as a diverging duct scramjet combustor. The facility operated at a total temperature of 1000 K and total pressure of 330 kPa. Hydrogen (H2), ethylene (C2H4) and methane (CH4) were used as fuels, and a wide range of global equivalence ratios were tested. The main data obtained were wall static pressure measurements, and the presence of combustion was determined based on the pressure rises obtained. Supersonic and dual-mode combustion were achieved with hydrogen and ethylene fuel, whereas very limited heat release was obtained with the methane. Global operability limits were determined to be 0.07 < Ï < 0.31 for hydrogen, and 0.14 < Ï < 0.48 for ethylene. The hydrogen fuel data for the aeroramp/torch system was compared to data from a physical 10 unswept compression ramp injector and similar performance was found with the two arrangements. With hydrogen and ethylene as fuels and the aeroramp/plasma-torch system, the effect of varying the air total temperature was investigated. Supersonic combustion was achieved with temperatures as low as 530K and 680K for the two fuels, respectively. These temperatures are facility/operational limits, not combustion limits. The pressure profiles were analyzed using the Ramjet Propulsion Analysis (RJPA) code. Results indicate that both supersonic and dual-mode ramjet combustion were achieved. Combustion efficiencies varied with Ï from a high of about 75% to a low of about 45% at the highest Ï . With a theoretical diffuser and nozzle assumed for the configuration and engine, thrust was computed for each fuel. Fuel specific impulse was on average 3000 and 1000 seconds for hydrogen and ethylene respectively, and air specific impulse varied from a low of about 9 sec to a high of about 24 sec (for both fuels) for the To = 1000K test condition. The GASP RANS code was used to numerically simulate the injection and mixing process of the fuels. The results of this study were very useful in determining the suitability of the selected plasma torch locations. Further, this tool can be used to determine whether combustion is theoretically possible or not. / Ph. D.

Hypersonic vehicles, based on scramjet engines, have the potential to deliver inexpensive access to space when compared with rocket propulsion. The technology, however, is in its infancy and there is still much to be learned from fundamental studies.¶ Flows that represent the conditions inside a scramjet engine can be generated in ground tests using a free-piston shock tunnel and a combustor model. These facilities provide a convenient location for fundamental studies and principles learned during ground tests can be applied to the design of a full-scale vehicle.¶ A wide range of diagnostics have been used for studying scramjet flows, including surface measurements and optical visualisation techniques.¶ The aim of this work is to test the effectiveness of tunable diode laser absorption spectroscopy (TDLAS) as a scramjet diagnostic.¶ TDLAS utilises the spectrally narrow emission from a diode laser to probe individual absorption lines of a target species. By varying the diode laser injection current, the laser emission wavelength can be scanned to rapidly obtain a profile of the spectral line. TDLAS has been used previously for gas-dynamic sensing applications and, in the configuration used in this work, is sensitive to temperature and water vapour concentration.¶ The design of the sensor was guided by previous work. It incorporated aspects of designs that were considered to be well suited to the present application. Aspects of the design which were guided by the literature included the laser emission wavelength, the use of fibre optics and the detector used. The laser emission wavelength was near 1390 nm to coincide with relatively strong water vapour transitions. This wavelength allowed the use of telecommunications optical fibre and components for light delivery. Detection used a dual-beam, noise cancelling detector.¶ The sensor was validated before deployment in a low-pressure test cell and a hydrogen–air flame. Temperature and water concentration measurements were verified to within 5% up to 1550 K. Verification accuracy was limited by non-uniformity along the beam path during flame measurements.¶ Measurements were made in a scramjet combustor operating in a flow generated by the T3 shock tunnel at the Australian National University. Within the scramjet combustor, hydrogen was injected into a flame-holding cavity and the sensor was operated downstream in the expanded, supersonic, post-combustion flow. The sensor was operated at a maximum repetition rate of 20 kHz and could resolve variation in temperature and water concentration over the 3ms running time of the facility.¶ Results were repeatable and the measurement uncertainty was smaller than the turbulent fluctuations in the flow. The scramjet was operated at two fuel-lean equivalence ratios and the sensor was able to show differences between the two operating conditions. In addition, vertical traversal of the sensor revealed variation in flow conditions across the scramjet duct.¶ The effectiveness of the diagnostic was tested by comparing results with those from other measurement techniques, in particular pressure and OH fluorescence measurements, as well as comparison with computational simulation.¶ Combustion was noted at both of the tested operating conditions in data from all three measurement techniques.¶ Computation simulation of the scramjet flow significantly under-predicted the water vapour concentration. The discrepancy between experiments and simulation was not apparent in either the pressure measurements or the OH fluorescence, but was clear in the diode laser results.¶ The diode laser sensor, therefore, was able to produce quantitative results which were useful for comparison with a CFD model of the scramjet and were complimentary to information provided by other diagnostics.

Inlet-injection was motivated by the possibility for skin-friction reduction in the combustion chamber of a flight style, three-dimensional, scramjet engine. High Mach number flight, where skin friction in the combustion chamber is a significant proportion of the overall drag, is the regime of interest for this type of reduction. This is a result of high Mach number supersonic flow within the combustion chamber, coupled with high densities due to the compression process. The flight condition of interest was chosen to be Mach 8.0 at an altitude of 30km. This choice was dictated by near-term flight-testing capabilities. The approach was to design an inlet with a reduced contraction ratio. This would produce a relatively low-density combustion-chamber flow, that would, in turn, lead to lower viscous drag. Due to low temperatures in the combustion chamber, as a result of the reduced compression, a novel method of ignition was required. This fluid-dynamic ignition technique made use of inlet injection together with flow non-uniformities generated by the inlet. The inlet chosen for this purpose was a rectangular-to-elliptical-shape-transition inlet or REST inlet. The focus of the investigation, was therefore, to determine the potential for performance improvement using inlet injection of fuel. The general approach to the investigation was experimental, using a scramjet model consisting of inlet, combustion chamber and a truncated nozzle. Flow-path thrust-potential was used as the primary performance parameter, where the term `thrust-potential' is used to indicate the lack of full expansion. A secondary performance metric was combustion efficiency, determined by matching one-dimensional analysis to experimental pressure distributions. In addition to inlet-injection, conventional injection into the combustion-chamber was tested as the performance baseline. Based on findings from these tests, two additional methods of injection were investigated both having a combination of inlet and combustion-chamber injection. The general findings showed that inlet injection, in comparison to combustion-chamber injection, produced an increase in performance in terms of thrust-potential and combustion efficiency for supersonic combustion. This occurred over a range of equivalence ratios up to 1.0. However, the maximum thrust developed by inlet injection was limited by engine unstart. In terms of the maximum thrust-potential, combustion-chamber injection exceeded that of inlet injection but significantly higher fuelling was required and poor combustion efficiency persisted. In order to offset the limit in thrust production due to unstart, an alternative fuelling method was implemented. This took the form of partial injection of the fuel in the combustion chamber in combination with inlet injection. An increase in thrust-potential and combustion efficiency as a result of increased fuel coverage in areas of the combustion chamber, which were fuel lean under inlet-injection. A thrust potential level similar to that of combustion-chamber injection was achieved with significantly higher combustion efficiency and consequently a lower fuelling level. This type of combined-injection is an attractive option for fuel delivery at the nominal flight condition. An additional finding for combustion-chamber and combined injection was that very high equivalence ratios led to separated flow in the combustion chamber and isolator. This was a result of excessive heat release producing an adverse pressure gradient in the engine. This mode of operation showed high levels of thrust-potential at equivalence ratios in excess of 1.0. Although interesting, these findings were outside the scope of the investigation since the flow within the combustion chamber is no longer purely supersonic.

One of the major challenges of scramjet propulsion remains the generation of sufficient thrust to overcome the large drag of hypersonic vehicles. Since the viscous drag constitutes a large portion of the overall drag, its mitigation offers potential for performance improvement. Viscous drag is generated on all wetted surfaces of the vehicle but is largest in the scramjet combustion chamber, where the fluid not only has a high flow speed, but also a high density. Reduction of the skin friction drag in the combustor hence promises large improvements to the efficiency of the propulsive system. Stalker (2005) proposed a novel approach to skin friction reduction that is based on the combustion of hydrogen within the turbulent boundary layer of supersonic or hypersonic flow. An extension to the theory of van Driest II was developed that suggests that the effectiveness of this method is significantly superior to that of film cooling without combustion effects. In essence, the combustion heat release reduces the velocity gradient at the wall and the density in the boundary layer so that the momentum transfer to the wall is decreased. This work investigates the applicability of this skin friction reduction method to scramjet combustors that would operate at flight Mach numbers between 8 and 13 at altitudes between 34 and 39 km. The corresponding combustor Mach number is approximately 4.5 and the total enthalpies are between 3.6 and 7.8 MJ/kg. Experiments that directly measured the skin friction drag on the internal scramjet combustor surface were conducted in the T4 Stalker tube at The University of Queensland. A constant area, axisymmetric combustor was tested with a matching constant area, axisymmetric inlet that did not compress the oncoming flow. Therefore, the experiments were of a quasi-direct-connect nature where the inlet was used to condition the wall layer of the flow that enters the combustion chamber. The start of the combustor was formed by a step at the end of the inlet which contained an annular slot for the injection of the gaseous hydrogen fuel. Fuel was injected tangentially to the main stream flow into the circular combustor as a uniform layer underneath the established boundary layer from this annular slot. Combustion was monitored via the measurement of the axial pressure distribution in the combustor and viscous forces on the combustor were measured with a stress wave force balance. Two different inlet lengths were tested to assess the effect of the boundary layer state and thickness on the ignition and combustion of the injected hydrogen. The leading edge of the inlet was either sharp or blunt to investigate the effect of the hot gas that is contained in an entropy layer that is generated by a blunt leading edge. Finally, the diameter of the duct was varied to ensure that the experimental data was not subject to duct scaling effects. The effect on skin friction of the combustion of fuel in the boundary layer was assessed directly by measurement as well as analytically with several prediction methods. The experimental data show reductions of skin friction drag of up to 77% when stable combustion was established. A thick, turbulent boundary layer results in ignition for lower enthalpy conditions than a thin, laminar layer. The blunted leading edge configuration creates conditions that results in ignition of the injected fuel at all tested flow enthalpies and when a sharp leading edge configuration does not. Analytical predictions of the skin friction drag are in close agreement with the experimental data for fuel-off, film cooling and boundary layer combustion cases. It is demonstrated that the characteristics of boundary layer combustion do not change when the duct diameter is increased and the hydrogen mass flow rate per unit circumferential length is kept constant.

Cette étude est consacrée à l’analyse des écoulements réactifs supersoniques cisailléset, plus particulièrement, des couches de mélange compressibles pouvant se développerdans les moteurs ramjet et scramjet. Des méthodes numériques appropriées ont été implémentéeset vérifiées pour aboutir au développement d’un code de calcul numériquemassivement parallèle, appelé CREAMS (compressible reactive multi-species solver). Cedernier a été spécialement conçu pour conduire des simulations numériques haute précision(simulations numériques directes ou DNS) de ce type d’écoulements. Une attentionparticulière a été portée à la description des termes de transport moléculaire et des termessources chimiques de façon à considérer la description physique la plus fidèle possible desmélanges des gaz réactifs à haute vitesse, au sein desquelles les temps caractéristiqueschimiques et de mélange aux petites échelles sont susceptibles d’être du même ordre degrandeur. Les simulations des couches de mélange bidimensionnelles et tridimensionnelles,inertes et réactives, confirment l’importance des effets associés à la compressibilité et autaux de dégagement de chaleur. Les résultats ainsi obtenus diffèrent en certains points deceux issus d’autres simulations qui introduisaient certaines hypothèses simplificatrices :développement temporel, emploi d’une chimie globale ou encore lois de transport simplifiées.En revanche, ils reproduisent certains tendances déjà observées dans un certainnombre d’études expérimentales conduites dans des conditions similaires. / This study is devoted to the analysis of supersonic reactive shear flows and, in particular,compressible mixing layers that can develop inside the ramjet and scramjet engines.Appropriate numerical methods have been implemented and tested to achieve the developmentof a massively parallel numerical solver, called CREAMS (compressible reactivemulti-species solver). This tool was designed to conduct high-precision numerical simulations(direct numerical simulations or DNS) of such flows. Particular attention waspaid to the description of the molecular transport terms and chemical source terms toconsider the most accurate physical description of reactive gas mixtures at high velocity,in which the chemical and mixing time scales, corresponding to the smallest scalesof the flow, are susceptible to be of the same order of magnitude. Simulations of twoandthree-dimensional, inert and reactive, mixing layers confirm the importance of theeffects associated with compressibility and rate of heat release. The results obtained differin some points from other simulations which introduced simplifying assumptions such astemporal development, use of a global chemistry or a simplified description of the moleculartransport terms. Nevertheless, they reproduce some trends already observed in severalexperimental studies conducted under similar conditions.