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Rowan, Scott A. "Viscous drag reduction in a scramjet combustor /". St. Lucia, Qld, 2003. http://www.library.uq.edu.au/pdfserve.php?image=thesisabs/absthe17438.pdf.
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Stouffer, Scott David. "The effect of flow structure on the combustion and heat transfer in a scramjet combustor". Diss., Virginia Tech, 1995. http://hdl.handle.net/10919/39116.
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Mundis, Nathan L. "Magnetohydrodynamic power generation in a scramjet using a post combustor generator". Diss., Rolla, Mo. : University of Missouri-Rolla, 2007. http://scholarsmine.umr.edu/thesis/pdf/Mundis_09007dcc8043ee98.pdf.
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Thesis (M.S.)--University of Missouri--Rolla, 2007.Vita. The entire thesis text is included in file. Title from title screen of thesis/dissertation PDF file (viewed March 25, 2008) Includes bibliographical references (p. 95-97).
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Corbin, Christopher Ryan. "Design and Analysis of a Mach 3 Dual Mode Scramjet Combustor". Wright State University / OhioLINK, 2008. http://rave.ohiolink.edu/etdc/view?acc_num=wright1208370076.
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Milligan, Ryan Timothy. "DUAL MODE SCRAMJET: A COMPUTATIONAL INVESTIGATION ON COMBUSTOR DESIGN AND OPERATION". Wright State University / OhioLINK, 2009. http://rave.ohiolink.edu/etdc/view?acc_num=wright1251725076.
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Malo-Molina, Faure Joel. "Numerical study of innovative scramjet inlets coupled to combustors using hydrocarbon-air mixture". Diss., Georgia Institute of Technology, 2010. http://hdl.handle.net/1853/33906.
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To advance the design of hypersonic vehicles, high-fidelity multi-physics CFD is used to characterize 3-D scramjet flow-fields in two novel streamline traced configurations. The two inlets, Jaws and Scoop, are analyzed and compared to a traditional rectangular inlet used as a baseline for on/off-design conditions. The flight trajectory conditions selected are Mach 6 and a dynamic pressure of 1,500 psf (71.82 kPa). Analysis of these hypersonic inlets is performed to investigate distortion effects downstream with multiple single cavity combustors acting as flame holders, and several fuel injection strategies. The best integrated scramjet inlet/combustor design is identified. The flow physics is investigated and the integrated performance impact of the two innovative scramjet inlet designs is quantified. Frozen and finite rate chemistry is simulated with 13 gaseous species and 20 reactions for an Ethylene/air finite-rate chemical model. In addition, URANS and LES modeling are compared to explore overall flow structure and to contrast individual numerical methods.
The flow distortion in Jaws and Scoop is similar to some of the distortion in the traditional rectangular inlet, despite design differences. The baseline and Jaws performance attributes are stronger than Scoop, but Jaws accomplishes this while eradicating the cowl lip interaction, and lessening the total drag and spillage penalties.
The innovative inlets work best on-design, whereas for off-design, the traditional inlet is best. Early pressure losses and flow distortions in the isolator aid the mixing of air and fuel, and improve the overall efficiency of the system. Although the trends observed with and without chemical reactions are similar, the former yields roughly 10% higher mixing efficiency and upstream reactions are present. These show a significant impact on downstream development. Unsteadiness in the combustor increases the mixing efficiency, varying the flame anchoring and combustion pressure effects upstream of the step.
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Griffiths, Alan David, i alan griffiths@anu edu au. "Development and demonstration of a diode laser sensor for a scramjet combustor". The Australian National University. Faculty of Science, 2005. http://thesis.anu.edu.au./public/adt-ANU20051114.132736.
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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 hydrogenair 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.
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Griffiths, Alan David. "Development and demonstration of a diode laser sensor for a scramjet combustor /". View thesis entry in Australian Digital Theses, 2005. http://thesis.anu.edu.au/public/adt-ANU20051114.132736/index.html.
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Etheridge, Steven J. "Effect of Flow Distortion on Fuel Mixing and Combustion in an Upstream-Fueled Cavity Flameholder for a Supersonic Combustor". University of Cincinnati / OhioLINK, 2012. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1353100774.
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McDaniel, Keith Scott. "Three Dimensional Simulation of Time-Dependent Scramjet Isolator /Combustor Flowfields Implemented on Parallel Architectures". NCSU, 2001. http://www.lib.ncsu.edu/theses/available/etd-20001228-204538.
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McDaniel, Keith S. Three Dimensional Simulation of Time-DependentScramjet Isolator / Combustor Flowfields Implemented onParallel Architectures, ( Under the directions of Dr. J. R. Edwards). The development of a parallel Navier-Stokes solver for computing time-dependent,three-dimensional reacting flowfields within scramjet (supersonic combusting ramjet)engines is presented in this work. The algorithm combines low-diffusion upwinding methods, timeaccurate implicit integration techniques, and domain decomposition strategies to yield an effectiveapproach for large-scale simulations. The algorithm is mapped to a distributed memoryIBM SP-2 architecture and a shared memory Compaq ES-40 architecture using the MPI-1 message-passingstandard. Two and three-dimensional simulations of time-dependent hydrogen fuel injection into a modelscramjet isolator / combustor configuration at two equivalence ratios are performed. Thesesimulations are used to gain knowledge of engine operability, inlet performance, isolatorperformance, fuel air mixing, flame holding, mode transition, and engine unstart.Results for an injection at a ratio of 0.29 show qualitative agreement withexperiment for the two-dimensional case, but revealed a slow progression towardengine unstart for the three-dimensional case. Injection at an equivalence ratio of 0.61resulted in engine unstart for both two-dimensional and three-dimensional cases.Engine unstart for the three-dimensional case occurs as a response to the formation and growthof large pockets of reversed flow along the combustor side wall. These structuresdevelop at an incipient pressure above 154 kPa and result in significant blockage of the core flow,additional compression, and chemical reaction within the boundary layer. All of these factors promotea much more rapid unstart as compared with the two-dimensional case.
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Ahuja, Vivek Hartfield Roy J. "Optimization of fuel-air mixing for a scramjet combustor geometry using CFD and a genetic algorithm". Auburn, Ala, 2008. http://hdl.handle.net/10415/1406.
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McGuire, Jeffrey Robert Aerospace Civil & Mechanical Engineering Australian Defence Force Academy UNSW. "Ignition enhancement for scramjet combustion". Awarded by:University of New South Wales - Australian Defence Force Academy. School of Aerospace, Civil and Mechanical Engineering, 2007. http://handle.unsw.edu.au/1959.4/38748.
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The process of shock-induced ignition has been investigated both computa- tionally and experimentally, with particular emphasis on the concept of radical farming. The first component of the investigation contained Computational Fluid Dynamic (CFD) calculations of an ignition delay study, a 2D pre-mixed flow over flat plate at a constant angle to the freestream, and through a generic 2D scramjet model. The focal point of the investigation however examined the complex 3D flow through a generic scramjet model. Five experimental test conditions were ex- amined over flow enthalpies from 3.4 MJ/kg to 6.4 MJ/kg. All test conditions simulated flight at 21000 metres ([symbol=almost equal to] 70000 ft), while the equivalent flight Mach number varied from approximately 8.5 at the lowest enthalpy, to approximately Mach 12 at the highest enthalpy condition. The presence of H2 fuel injected in the intake caused a separated region to form on the lower surface of the model at the entrance to the combustor. A fraction of the total mass of fuel was entrained in this separated region, providing long residence times, hence increased time for the chemical reactions that lead to ignition to occur. In addition, extremely high temperatures were found to exist between each fuel jet. Both fuel and air are present in these regions, therefore the chance of ignition in these regions is high. Streamlines passing through the recirculation zone ignited within this zone, while streamlines passing between the fuel jets ignited soon after entry into the combustor. The first instance of a pressure rise from combustion was observed on the centreline of the model where the reflected bow shock around the fuel jets crossed the centreline of the combus- tor. Upstream of this location the static pressure of the flow was too low for the chemical reactions that release heat to occur. The comparison between the experimental and computational results was lim- ited due to inaccuracies in modelling the thermal state of the gas in the CFD calculations. The gas was modelled as being in a state of thermal equilibrium at all times, which incorrectly models the freestream flow from the nozzle of the shock tunnel, and also the flow downstream of oblique shock wave within the scramjet model. As a result combustion occurs sooner in the CFD calculations than in the experimental result.
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Hoste, Jimmy-John O. E. "Scramjet combustion modeling using eddy dissipation model". Thesis, University of Strathclyde, 2018. http://digitool.lib.strath.ac.uk:80/R/?func=dbin-jump-full&object_id=30307.
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In order to aid in the design of scramjet propulsion systems at high Mach number operation, this works considers the Eddy Dissipation Model (EDM) to describe the combustion process inside an open-access Computational Fluid Dynamics (CFD) solver. Typical CFD modeling approaches for turbulent supersonic reacting flows are associated with a high computational cost. This in turn inhibits the use of CFD in scramjet combustor design or in higher level preliminary designs such as the trajectory optimization process of a scramjet powered vehicle. Instead, low-fidelity models are preferred to charaterize the propulsion system in the latter type of application. The EDM relies on simplified assumptions regarding the combustion process whose validity is thought to be prevalent at high Mach number scramjet operation. It is therefore a suitable candidate model in order to introduce more routinely CFD in scramjet preliminary design phases. As part of the present work, first steps include the selection of an open-source CFD solver followed by several validation studies. After its implementation, a critical numerical analysis of the EDM is performed by considering three hydrogen-fueled experimental scramjet configurations with different fuel injection approaches. Its application is further investigated with a mainly kinetically controlled scramjet design where the underlying assumptions of the EDM are not valid anymore. Finally, the EDM is applied to a combustor design problem demonstrating the metrics of interest that can be relied on for this task.
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Odam, Judy. "Scramjet experiments using radical farming /". [St. Lucia, Qld.], 2004. http://adt.library.uq.edu.au/public/adt-QU20041206.101729/index.html.
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Prebola, John L. Jr. "Performance of a Plasma Torch with Hydrocarbon Feedstocks for Use in Scramjet Combustion". Thesis, Virginia Tech, 1998. http://hdl.handle.net/10919/36941.
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Research was conducted at Virginia Tech on a high-pressure uncooled plasma torch to study torch operational characteristics with hydrocarbon feedstocks and to determine the feasibility of using the torch as an igniter in scramjet applications. Operational characteristics studied included electrical properties, such as arc stability, voltage-current characteristics and start/re-start capabilities, and mechanical properties, such as coking, electrode erosion and transient to steady-state torch body temperature trends. Possible use of the plasma torch as an igniter in high-speed combustion environments was investigated through the use of emission spectroscopy and a NASA chemical kinetics code.
All feedstocks tested; argon, methane, ethylene and propylene, were able to start. The voltage data indicated that there were two preferred operating modes, which were well defined for methane. For all gases, a higher current setting, on the order of 40 A, led to more stable torch operation. A low intensity, high frequency current applied to the torch, along with the primary DC current, resulted in virtual elimination of soot deposits on the anodes. Electrode erosion was found to multiply each time the complexity of the hydrocarbon was increased. Audio and high-speed visual analysis led to identification of 180 Hz plasma formation cycle, related to the three-phase power supply. The spectroscopic analysis aided in the identification of combustion enhancing radicals being produced by the torch, and results of the chemical kinetics analysis verified combustion enhancement and radical production through the use of a basic plasma model. Overall, the results of this study indicate that the plasma torch is a promising source for scramjet ignition, and further study is warranted.Master of Science
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Rock, Christopher. "Experimental Studies of Injector Array Configurations for Circular Scramjet Combustors". Diss., Virginia Tech, 2010. http://hdl.handle.net/10919/77208.
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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.
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MURUGAPPAN, SHANMUGAM. "INNOVATIVE TECHNIQUES TO IMPROVE MIXING AND PENETRATION IN SCRAMJET COMBUSTORS". University of Cincinnati / OhioLINK, 2005. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1109697512.
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Miki, Kenji. "Simulation of magnetohydrodynamics turbulence with application to plasma-assisted supersonic combustion". Diss., Atlanta, Ga. : Georgia Institute of Technology, 2009. http://hdl.handle.net/1853/26605.
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Thesis (Ph.D)--Aerospace Engineering, Georgia Institute of Technology, 2009.Committee Chair: Menon Suresh; Committee Co-Chair: Jagoda Jeff; Committee Member: Ruffin Stephen; Committee Member: Thorsten Stoesser; Committee Member: Walker Mitchell. Part of the SMARTech Electronic Thesis and Dissertation Collection.
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Cocks, Peter. "Large eddy simulation of supersonic combustion with application to scramjet engines". Thesis, University of Cambridge, 2011. https://www.repository.cam.ac.uk/handle/1810/239344.
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This work evaluates the capabilities of the RANS and LES techniques for the simulation of high speed reacting flows. These methods are used to gain further insight into the physics encountered and regimes present in supersonic combustion. The target application of this research is the scramjet engine, a propulsion system of great promise for efficient hypersonic flight. In order to conduct this work a new highly parallelised code, PULSAR, is developed. PULSAR is capable of simulating complex chemistry combustion in highly compressible flows, based on a second order upwind method to provide a monotonic solution in the presence of high gradient physics. Through the simulation of a non-reacting supersonic coaxial helium jet the RANS method is shown to be sensitive to constants involved in the modelling process. The LES technique is more computationally demanding but is shown to be much less sensitive to these model parameters. Nevertheless, LES results are shown to be sensitive to the nature of turbulence at the inflow; however this information can be experimentally obtained. The SCHOLAR test case is used to validate the reacting aspects of PULSAR. Comparing RANS results from laminar chemistry and assumed PDF combustion model simulations, the influence of turbulence-chemistry interactions in supersonic combustion is shown to be small. In the presence of reactions, the RANS results are sensitive to inflow turbulence, due to its influence on mixing. From complex chemistry simulations the combustion behaviour is evaluated to sit between the flamelet and distributed reaction regimes. LES results allow an evaluation of the physics involved, with a pair of coherent vortices identified as the dominant influence on mixing for the oblique wall fuel injection method. It is shown that inflow turbulence has a significant impact on the behaviour of these vortices and hence it is vital for turbulence intensities and length scales to be measured by experimentalists, in order for accurate simulations to be possible.
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Eugênio, Ribeiro Fábio Henrique. "Numerical Simulation of Turbulent Combustion in Situations Relevant to Scramjet Engine Propulsion". Thesis, Chasseneuil-du-Poitou, Ecole nationale supérieure de mécanique et d'aérotechnique, 2019. http://www.theses.fr/2019ESMA0001/document.
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Les super-statoréacteurs sont des systèmes de propulsion aérobie à grande vitesse qui ne nécessitent pas d’éléments rotatifs pour comprimer l’écoulement d’air. Celui-ci est comprimé dynamiquement par un système d’admission intégré dans le véhicule, atteignant la pression et la température requises pour que la combustion puisse s’opérer dans la chambre de combustion. La chambre de combustion est traversée par un écoulement supersonique dans ce type de moteur, ce qui limite considérablement le temps disponible pour injecter le carburant, le mélanger avec un oxydant, enflammer le mélange obtenu et parvenir à une combustion complète. Les cavités peuvent être utilisées pour augmenter le temps de séjour sans perte excessive de pression totale et sont donc utilisées comme éléments de stabilisation dans les chambres de combustion supersonique. Cette thèse se concentre sur l’étude du mécanisme de stabilisation et des interactions chimie-turbulence dans le cas d’une injection pariétale de combustible dans un écoulement supersonique d’air vicié en amont d’une cavité carrée. Les conditions d’écoulement réactif à grande vitesse correspondantes sont examinées sur la base de simulations numériques d’un modèle de scramjet représentatif d’expériences effectuées précédemment à l’Université du Michigan. Les calculs sont effectués avec le solveur CREAMS, développé pour effectuer la simulation numérique d’écoulements multi-espèces réactifs compressibles sur des architectures massivement parallèles. Le solveur utilise des schémas numériques d’ordre élevé appliqués sur des maillages structurées et la géométrie de la chambre de combustion est modélisée à l’aide d’une méthode de frontières immergées (IBM). Les simulations LES font usage du modèle wall-adapting local eddy (WALE). Deux températures distinctes sont considérées dans l’écoulement entrant d’air vicié pour étudier la stabilisation de la combustion.Une attention particulière est accordée à l’analyse de la topologie et de la structure des écoulements réactifs, les régimes de combustion sont analysés sur la base de diagrammes standards de combustion turbulenteScramjet engines are high-speed air breathing propulsion systems that do not require rotating elements to compress the air inlet stream. The flow is compressed dynamically through a supersonic intake system integrated in the aircraft’s forebody, reaching the required pressure and temperature for combustion to proceed within the combustor in this kind of engine. The combustion chamber is crossed by a supersonic flow, which limits severely the time available to inject fuel, mix it with oxidizer, ignite the resulting mixture and reach complete combustion. Cavities can be used to increase the residence time without excessive total pressure loss and are therefore used as flame holders in supersonic combustors.This thesis focuses in studying the flame stabilization mechanism and turbulence-chemistry interactions for a jet in a supersonic crossflow (JISCF) of vitiated air with hydrogen injection upstream of a wall-mounted squared cavity. The corresponding reactive high-speed flow conditions are scrutinized on the basis of numerical simulations of a scramjet model representative of experiments previously conducted at the University of Michigan. The computations are performed with the high-performance computational solver CREAMS, developed to perform the numerical simulation of compressible reactive multi-component flows on massively-parallel architectures. The solver makes use of high-order precision numerical schemes applied on structured meshes and the combustion chamber geometry is modeled by using the Immersed Boundary Method (IBM) algorithm. The present set of computations is conducted within the LES framework and the subgrid viscosity is treated with the wall-adapting local eddy (WALE)model. Two distinct temperatures are considered in the inlet vitiated airstream to study combustion stabilization. Special emphasis is placed on the analysis of the reactive flow topology and structure,and the combustion regimes are analyzed on the basis of standard turbulent combustion diagrams
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Karl, Sebastian. "Numerical Investigation of a Generic Scramjet Configuration". Doctoral thesis, Saechsische Landesbibliothek- Staats- und Universitaetsbibliothek Dresden, 2011. http://nbn-resolving.de/urn:nbn:de:bsz:14-qucosa-68695.
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A Supersonic Combustion Ramjet (scramjet) is, at least in theory, an efficient air-breathing propulsion system for sustained hypersonic flight at Mach numbers above approximately M=5. Important design issues for such hypersonic propulsion systems, are the lack of ground based facilities capable of testing a full-sized engine at cruise flight conditions and the absence of general scaling laws for the extrapolation of wind tunnel data to flight configurations. Therefore, there is a strong need for the development and validation of CFD tools to support the design process of scramjet-powered vehicles. The aims of this thesis are, in this context, to assess the applicability of, to further develop, and to validate the DLR TAU flow solver for the CFD analysis of the complete flow-path of a scramjet vehicle. The basis of this validation and of the identification of critical modelling assumptions is the recalculation of a series of wind tunnel tests of the HyShot II generic scramjet configuration that were performed in the High Enthalpy Shock Tunnel Göttingen (HEG) at the German Aerospace Center, DLRStaustrahlantriebe, bei denen sich die Strömung im gesamten Triebwerksbereich im Überschall befindet (supersonic combustion ramjets, Scramjets), stellen ein - zumindest theoretisch - effektives Antriebessystem für den Hyperschallflug im Machzahlbereich von M > 5 dar. Die Auslegung und der Entwurf von luftatmenden Hyperschallantrieben sind in der Praxis mit Schwierigkeiten verbunden. Der Einsatz von Bodenversuchsanlagen ist auf kleinskalige Konfigurationen oder einzelne Triebwerkskomponenten begrenzt. Die Ergebnisse von numerischen Strömungssimulationsverfahren sind mit hohen Unsicherheiten behaftet, die ihren Ursprung in der Modellbildung für die komplexen Strömungsphänomene in chemisch reagierenden, kompressiblen und turbulenten Über- und Hyperschallströmungen haben. Weiterhin existieren keine allgemein gültigen Skalierungsgesetze um Aussagen aus Windkanalexperimenten auf Flugkonfigurationen zu übertragen.Die vorliegende Arbeit beschäftigt sich in diesem Zusammenhang mit der Erweiterung des DLRStrömungslösers TAU für die Berechnung von Überschallverbrennungsphänomenen in Scramjets sowie mit der Anwendung des Verfahrens für die numerische Analyse von Windkanalexperimenten, die im Hochenthalpiekanal Göttingen (HEG) des Deutschen Zentrums für Luft- und Raumfahrt (DLR) zur Untersuchung der generischen HyShot II Scramjet-Konfiguration durchgeführt wurden. Die wichtigsten Ziele waren die genaue Charakterisierung der freien Anströmung im Windkanal, der Nachweis der Anwendbarkeit des verwendeten Rechenverfahrens und die Analyse des Einflusses verschiedener numerischer Modellierungsansätze für die Strömungssimulation in Scramjets sowie die Nutzung der numerischen Daten für eine verbesserte Interpretation der experimentellen Ergebnisse
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Ruan, Jiangheng Loïc. "Large eddy simulation of supersonic combustion in cavity-based scramjets". Thesis, Normandie, 2019. http://www.theses.fr/2019NORMIR14.
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Les dernières décennies ont été marquées par la course aux technologies hypersoniques. Voler à une vitesse hypersonique pourrait être possible avec les superstatoréacteurs. Mais le principal problème de ce moteur est le court temps de résidence du combustible dans la chambre de combustion, qui est de l'ordre de la milliseconde, rendant le mélange et la combustion difficile. L'ajout d'une cavité dans les superstatoréacteurs pourrait palier à ce problème grâce aux zones de recirculation de la cavité qui emprisonnent les gaz brulés, et permettent ainsi de rallumer continuellement le combustible. Grâce à l'essor de l'informatique, une simulation aux grandes échelles d'un telle configuration devient possible de nos jours. Les objectives de la thèse sont dans un premier temps d'évaluer la capacité d'une simulation aux grandes échelles à prédire des écoulements compressibles réactifs, et dans un second temps, de comprendre les phénomènes propres aux superstatoréacteurs à cavitéThe last decades have been marked by great progress in hypersonic technologies. The scramjet seems to be able to cope with these hypersonic speeds even today. The main problem to overcome is the short residence time of the fuel in the combustion chamber. This time being of the order of a millisecond, mixing and combustion cannot operate efficiently making the flameholding a challenging task. The cavity-based scramjets have been considered as a promising solution because the recirculation of the combustion gases inside of it makes it possible to ignite the reaction mixture continuously. Due to the increase in high performance computing, the use of Large-Eddy Simulation for supersonic combustion is now becoming relevant. The objectives of the present study are twofold: first, assess the ability of the LES technique to predict compressible multi-species reacting flows; and second, provide some fundamental aspects of cavity-based scramjet
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Zang, Andrew Henry. "Fuel injection in scramjets mixing enhancement and combustion characterization experiments /". College Park, Md. : University of Maryland, 2005. http://hdl.handle.net/1903/2559.
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Thesis (M.S.) -- University of Maryland, College Park, 2005.Thesis research directed by: Dept. of Aerospace Engineering. Title from t.p. of PDF. Includes bibliographical references. Published by UMI Dissertation Services, Ann Arbor, Mich. Also available in paper.
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Dröske, Nils [Verfasser]. "Investigation of Thermal Loads onto a Cooled Strut Injector inside a Scramjet Combustion Chamber / Nils Dröske". München : Verlag Dr. Hut, 2016. http://d-nb.info/1120763681/34.
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Dröske, Nils Christoph [Verfasser]. "Investigation of Thermal Loads onto a Cooled Strut Injector inside a Scramjet Combustion Chamber / Nils Dröske". München : Verlag Dr. Hut, 2016. http://d-nb.info/1120763681/34.
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Del, Rio Francesco. "Distortion mechanism in supersonic combustion ramjet engines". Master's thesis, Alma Mater Studiorum - Università di Bologna, 2018.
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Il mio lavoro di tesi è stato incentrato sulla progettazione e la realizzazione di un prototipo di isolator (componente necessaria per il funzionamento dei motori scramjet, utilizzati per velivoli aerospaziali ipersonici) in grado di generare tramite un opportuno dispositivo il meccanismo fluidodinamico che in letteratura viene definito "distortion mechanism". Tramite la tecnica fotografica denominata Schlieren, la quale sfrutta i gradienti di densità all’interno del fluido in esame, ho fotografato le onde di shock generate dal meccanismo suddetto, rendendo così possibile la comprensione del comportamento di queste onde e delle loro interazioni con il boundary layer, con le pareti, ma soprattutto dell’influenza che esse hanno sulle prestazioni di un eventuale propulsore. Da qui è partita una analisi sulle interazioni shock-shock e shock-boundary layer: quest’ultimo fenomeno è di grande interesse in quanto si è notato che non solo viene attivato un meccanismo di distorsione dell’onda stessa, ma che addirittura si manifesta la separazione dello strato limite, generando complessi fenomeni fluidodinamici e termodinamici i quali decrementano l’efficienza non solo dell’isolator bensì del motore stesso.È stato infine previsto come le onde di shock che si propagavano nell’isolator avrebbero potuto affliggere il mixing e la combustione nell’ultimo stage del prototipo, evidenziando le conseguenze che avrebbero generato sull’efficienza generale del ciclo termodinamico.
Per concludere il mio lavoro di tesi ho sviluppato alcuni tools in ambiente Matlab utili per poter calcolare le proprietà termodinamiche di un fluido che entra in un inlet di uno scramjet. Per motivi di complessità del problema e per la non assoluta certezza dei fenomeni fluidodinamici e termodinamici che realmente accadono in questi motori (in 3-D), le equazioni utilizzate all’interno del codice sono utili per un’analisi di un fluido quasi monodimensionale.
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Campioli, Theresa Lynn. "Computational Studies of Penetration and Mixing for Complex Jet Injectors to Aid in Design of Hypersonic Systems". Diss., Virginia Tech, 2007. http://hdl.handle.net/10919/28132.
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A computational study of sonic light-gas jet injection into a supersonic cross flow was conducted. The scope of the numerical analysis encompassed many studies that affect how the flow-field is numerically modeled and the behavior, specifically mixing, of the flow-field itself. A single, round injector was used for the Baseline design. Simulated conditions involved sonic injection of helium heated to 313 K into a Mach 4 air cross-stream with average Reynolds number 5.77 e+7 per meter and a freestream momentum flux ratio of 2.1. Experiments at these conditions were available for comparison. The primary numerical flow solver employed was GASP v. 4.2. The Menter Shear Stress Transport (SST) turbulence model was used, since the algorithm has good capability of solving both wall-bounded and free-shear flows. The SST model was able to capture the mixing behavior of the complex flow-field. Important numerical parameters that affect the capabilities of the numerical solver were studied for the Baseline injector. These sensitivity studies varied the choice of turbulent Prandtl number, Schmidt number, freestream turbulence intensity, boundary layer size, steady and unsteady approaches and computational software packages. A decrease in the turbulent Prandtl number resulted in better mixing behavior of the prediction and better agreement with the experiment. An increase in the turbulent Schmidt number had a small adverse effect on the predictions. The mixing characteristics remained constant with an increase in freestream turbulence intensity. The best Baseline prediction was then compared to three different injector configurations: an aerodynamic ramp consisting of four injectors in an array, a diamond injector both aligned and yawed 15° to the oncoming flow. The Computational Fluid Dynamics (CFD) tools were more accurate compared to experiment in the prediction of the aeroramp injector than the diamond-shaped injectors. The aeroramp injector slightly improved mixing efficiency over the Baseline injector at these conditions. Both of the diamond-shaped injectors had similar mixing as the Baseline injector but did not predict significant improvement in penetration for the analyzed conditions. Additional studies involving the interaction of transverse injection with impinging oblique shock waves were performed. The impingement of a shock upon light gas jet injection increased mixing. The closer the shock is to the injection point, the larger the effect on mixing and vorticity. The last analyses involved a numerical comparison of a non-reacting model to a reacting hydrogen-air model. The reacting analysis prediction had an improved spreading rate and larger counter-rotating vortex pair with downstream distance over the non-reacting analysis. The mixing was not significantly altered by the addition of hydrogen-air reactions to the numerical equations. The numerical tools used are capable of reasonable accuracy in predicting the complex flow-field of jet injection into a supersonic freestream with proper choice of models and parameters. Numerical modeling offers a way to study the entire flow-field thoroughly in a cost and time efficient manner.Ph. D.
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Gallimore, Scott D. Jr. "Operation of a High-Pressure Uncooled Plasma Torch with Hydrocarbon Feedstocks". Thesis, Virginia Tech, 1998. http://hdl.handle.net/10919/36917.
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The main scope of this project was to determine if a plasma torch could operate on pure hydrocarbon feedstocks and, if so, to catalogue the torch operational characteristics. The future goal of the project is to design a plasma torch for supersonic combustion applications that operates off of the vehicle main fuel supply to simplify onboard fuel systems. Experiments were conducted with argon, methane, ethylene and propylene. Spectrographic tests and tests designed to catalogue current/voltage characteristics, plasma jet phenomena, arc stability dependencies, electrode erosion rate and torch body temperature were performed.
Spectrographic analysis of the plasma jet exhaust confirmed the presence of combustion-enhancing radicals for each hydrocarbon gas tested. Also, it was discovered that simple hydrocarbon gases, such as methane, produced smooth torch operation, while even slightly more complex gases, ethylene and propylene, caused unsteady performance. Plasma jet oscillation was found to be related to the voltage waveform of the power supplies, indicating that plasma jet length and oscillation rate could be controlled by changing the input voltage.
The plasma torch for this study was proven to have the capability of operating with pure hydrocarbon feedstocks and producing radicals that are known to reduce combustion reaction rate times. The torch was demonstrated to have potential for use in supersonic combustion applications.Master of Science
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Bonanos, Aristides Michael. "Scramjet Operability Range Studies of an Integrated Aerodynamic-Ramp-Injector/Plasma-Torch Igniter with Hydrogen and Hydrocarbon Fuels". Diss., Virginia Tech, 2005. http://hdl.handle.net/10919/28847.
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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 10º 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.
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Dröske, Nils Christoph [Verfasser], i Jens von [Akademischer Betreuer] Wolfersdorf. "Investigation of thermal loads onto a cooled strut injector inside a scramjet combustion chamber / Nils Christoph Dröske ; Betreuer: Jens von Wolfersdorf". Stuttgart : Universitätsbibliothek der Universität Stuttgart, 2016. http://d-nb.info/1123572410/34.
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Axdahl, Erik Lee. "A study of premixed, shock-induced combustion with application to hypervelocity flight". Diss., Georgia Institute of Technology, 2013. http://hdl.handle.net/1853/50290.
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One of the current goals of research in hypersonic, airbreathing propulsion is access to higher Mach numbers. A strong driver of this goal is the desire to integrate a scramjet engine into a transatmospheric vehicle airframe in order to improve performance to low Earth orbit (LEO) or the performance of a semi-global transport. An engine concept designed to access hypervelocity speeds in excess of Mach 10 is the shock-induced combustion ramjet (i.e. shcramjet). This dissertation presents numerical studies simulating the physics of a shcramjet vehicle traveling at hypervelocity speeds with the goal of understanding the physics of fuel injection, wall autoignition mitigation, and combustion instability in this flow regime.
This research presents several unique contributions to the literature. First, different classes of injection are compared at the same flow conditions to evaluate their suitability for forebody injection. A novel comparison methodology is presented that allows for a technically defensible means of identifying outperforming concepts. Second, potential wall cooling schemes are identified and simulated in a parametric manner in order to identify promising autoignition mitigation methods. Finally, the presence of instabilities in the shock-induced combustion zone of the flowpath are assessed and the analysis of fundamental physics of blunt-body premixed, shock-induced combustion is accelerated through the reformulation of the Navier Stokes equations into a rapid analysis framework. The usefulness of such a framework for conducting parametric studies is demonstrated.
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Bouheraoua, Lisa. "Simulation aux grandes échelles et modélisation de la combustion supersonique". Thesis, Rouen, INSA, 2014. http://www.theses.fr/2014ISAM0022/document.
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Le travail de cette thèse est consacré à la simulation aux grandes échelles (LES) et à la modélisationde la combustion supersonique, dont l’application est rencontrée dans les moteurs detype scramjet. Dans ce contexte, une étude LES appliquée au cas d’une flamme supersoniquehydrogène-air (flamme de Cheng) a été effectuée sur trois niveaux de raffinements de maillage.Les résultats en termes de profils moyens et fluctuations de composition et de température sontconfrontés aux mesures expérimentales, et l’impact du raffinement de maillage est établi. Parailleurs, à partir des données issues de cette étude LES, une modélisation de la combustionturbulente dans un milieu fortement compressible est proposée sur la base d’une approche tabuléede la chimie. Une analyse temporelle des interactions choc/flamme a ensuite été menée,permettant de mettre en évidence la présence de structures transitoires ayant une influence surles processus de stabilisation de la flammeThis PhD study is focused on the large eddy simulation (LES) and on the modelisation of supersonic combustion as encountered in scramjet types engines. In this context, a LES study was performed for an hydrogen-air supersonic flame (Cheng’s flame) with three mesh refinement levels. The results obtained for mean and fluctuations of composition and temperature are compared to experimental measurements, and the impact of the grid resolution is established. Moreover, a modelisation of turbulent combustion in highly compressible flows is proposed based of tabulated chemistry approach. An analysis of the dynamics of shock/flame interaction was then conducted, and the presence of transient structures, which impact the flame stabilisation processes, was emphasized
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Perkins, Hugh Douglas. "Development and Demonstration of a Computational Tool for the Analysis of Particle Vitiation Effects in Hypersonic Propulsion Test Facilities". Case Western Reserve University School of Graduate Studies / OhioLINK, 2009. http://rave.ohiolink.edu/etdc/view?acc_num=case1227553721.
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Retaureau, Ghislain J. "On recessed cavity flame-holders in supersonic cross-flows". Diss., Georgia Institute of Technology, 2012. http://hdl.handle.net/1853/43703.
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Flame-holding in a recessed cavity is investigated experimentally in a Mach 2.5 preheated cross-flow for both stable and unstable combustion, with a relatively low preheating. Self-sustained combustion is investigated for stagnation pressures and temperatures reaching 1.4 MPa and 750 K. In particular, cavity blowout is characterized with respect to cavity aspect ratio (L/D =2.84 - 3.84), injection strategy (floor - ramp), aft ramp angle (90 deg - 22.5 deg) and multi-fuel mixture (CH₄-H₂ or CH₄-C₂H₄ blends). The results show that small hydrogen addition to methane leads to significant increase in flame stability, whereas ethylene addition has a more gradual effect. Since the multi-fuels used here are composed of a slow and a fast chemistry fuel, the resulting blowout region has a slow (methane dominant) and a fast (hydrogen or ethylene dominant) branch. Regardless of the fuel composition, the pressure at blowout is close to the non-reacting pressure imposed by the cross-flow, suggesting that combustion becomes potentially unsustainable in the cavity at the sub-atmospheric pressures encountered in these supersonic studies. The effect of preheating is also investigated and results show that the stability domain broadens with increasing stagnation temperature. However, smaller cavities appear less sensitive to the cross-flow preheating, and stable combustion is achieved over a smaller range of fuel flow rate, which may be the result of limited residence and mixing time. The blowout data point obtained at lower fuel flow rate fairly matches the empirical model developed by Rasmussen et al. for floor injection phi = 0.0028 Da^-.8, where phi is the equivalence ratio and Da the Damkohler number. An alternate model is proposed here that takes into account the ignition to scale the blowout data. Since the mass of air entrained into the cavity cannot be accurately estimated and the cavity temperature is only approximated from the wall temperature, the proposed scaling has some uncertainty. Nevertheless the new phi-Da scaling is shown to preserve the subtleties of the blowout trends as seen in the current experimental data.
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Shih, Ming Sam, i 施明憲. "Simulation of Shock Induced Mixing and Combustion Analysis on Scramjet Combustor". Thesis, 1995. http://ndltd.ncl.edu.tw/handle/64928395065623884774.
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謝宗燁. "Study of Combustion Effects for Hydrogen Injection in Scramjet Engine Combustor". Thesis, 2015. http://ndltd.ncl.edu.tw/handle/89420493458057462392.
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碩士逢甲大學
航太與系統工程學系
103
This study describes using the finite volume method to solve Reynold average Navier-Stokes Equations, in order to simulates the flow field of external intake compression ramp to the internal combustion chamber on a supersonic combustion ramjet engine(Scramjet Engine). And this thesis uses non-premixed combustion model to simulate the combustion reaction process of supersonic combustion. The operation theory of supersonic combustion ramjet engine is that in hypersonic flight conditions, air compressed by shock waves pass into the combustion chamber and produce combustion reaction with fuel. The flow through nozzle exchange for thrust at last. This research focuses on the phenomena of heat flow in the combustion chamber when the fluid combusts in the supersonic conditions. Previous study observes the sequence of external shock wave development through two dimension unsteady simulations with non-fuel injection. After that, the research observes the influence of Hydrogen injection velocity interact the combustion field by maintaining the total pressure and total temperature of fuel and changing the injection mach number to 2.5, 2.75, 3.0. This study discovers injection speed influence the deflection angle of shock wave. It indirectly affects times of shock reflection and merged position on the wall. The injection velocity also affects the thickness of the flame. Hydrogen injects more faster, the flame thickness become thinner. The research further explores a three-dimensional combustion chamber, and uses unsteady simulation to observe the shape of flame in the three-dimensional supersonic combustion flow field. Then we acquire the combustion phenomena of profile section through the presentation of streamlines for the combustion chamber along the axial channel. This research discovers the supersonic combustion flow patterns closely linked with shock waves reflection and intersection.
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Hwang, Kuh-Wen, i 黃顧文. "Cold & Hot Flow Analysis of SCRAMJET Combustor". Thesis, 1994. http://ndltd.ncl.edu.tw/handle/31460670763536126525.
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碩士國立中央大學
機械工程研究所
82
The two-dimensional full Navier-Stokes equations and species transport equations are solved numerically by employing the finite-volume, explicit MacCormack scheme. The two-layer algebraic turbulent model and one-step instantaneous chemical reaction are adopted to simulated the complicated turbulent reaction flow. First, the cold flows with fuel injections are solved to show the patterns of mixing and the recirculation zone in primarily determining the better location of fuel inlet which will be applied later for hot flow calculations. Subsequently, the reacting flows are further solved with the fuel injection locations determined by the cold flow mixing analysis. It shows that the fuel injected from the midpoint between the dump step plate and reattachment point will result in better thrust.
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Yeung, Moon-Tai. "Chemical Reactions in a Scramjet Combustor and Two-Dimensional Nozzles". Thesis, 1993. https://thesis.library.caltech.edu/6333/1/Yeung_mt_1993.pdf.
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Finite-rate chemistry of hydrogen-air combustion is to be investigated numerically in a one-dimensional constant pressure SCRAMJET combustor and two-dimensional nozzles. Detailed reaction mechanisms and temperature dependent thermodynamics are to be used in the models. The aspects of interest include the combustion characteristics at different fuel-air ratios, pressures and initial temperatures in the combustor. Methods for enhancing the combustion rate in the combustor is to be studied also. The effect of expansion rate on the hydrogen-air reactions is the prime focus of the nozzle calculation. The results from different inlet conditions and wall geometries are to be analyzed.
A computer model for a one-dimensional (channel-flow) combustor is constructed based on the chemical kinetics subroutine library CHEMKIN. Subsequent calculations show that the initial temperature is the most important parameter in the combustor. It is further discovered that certain reaction steps are responsible for the initial delay exhibited in all hydrogen-air combustion processes. Low temperature behavior is studied extensively and augmentation methods are developed. The introduction of a small percentage of the hydrogen radical into the initial mixture is found to be the most effective in reducing the reaction delay. The combustor pressure enters the overall reaction process in a linear manner. The calculations over five combustor pressures show that the initial delay in hydrogen-air reaction and the following period of explosion are proportional to the combustor pressure raised to certain powers.
The nozzle model is two-dimensional, steady and inviscid with no conductivity and diffusivity. Two schemes are developed to handle the boundary conditions. One is based on pure numerical interpolation/extrapolation methods while the other imposes analytical supersonic characteristic equations. The former scheme is found to be more efficient while the latter is more accurate. In analysing the response of the combustion product to an expansion, it is found that the formation of water is favoured by an expansion. A closer examination reveals that the behavior can be attributed to the abundance of free radicals in the nozzle inlet composition. Freezing is not clearly observed except for the NOx species.
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Griffiths, Alan David. "Development and demonstration of a diode laser sensor for a scramjet combustor". Phd thesis, 2005. http://hdl.handle.net/1885/47106.
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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.¶ ...
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Perez, Jaime Enrique. "EVALUATION OF GEOMETRIC SCALE EFFECTS FOR SCRAMJET ISOLATORS". 2010. http://trace.tennessee.edu/utk_gradthes/739.
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A numerical analysis was conducted to study the effects of geometrically scaling scramjet inlet-combustor isolators. Three-dimensional fully viscous numerical simulation of the flow inside constant area rectangular ducts, with a downstream back pressure condition, was analyzed using the SolidWorks Flow Simulation software. The baseline, or 1X, isolator configuration has a 1” x 2.67” cross section and 20” length. This baseline configuration was scaled up based on the 1X configuration mass flow to 10X and 100X configurations, with ten and one hundred times the mass flow rate, respectively. The isolator aspect ratio of 2.67 was held constant for all configurations. To provide for code validation, the Flow Simulation program was first used to analyze a converging-diverging channel and a wind tunnel nozzle. The channel case was compared with analytical theory and showed good agreement. The nozzle case was compared with AFRL experimental data and showed good agreement with the entrance and exit conditions (Pi0= 40 psia, Ti0= 530ºR, Pe= 18.86 psia, Te= 456ºR, respectively). While the boundary layer thickness remained constant, the boundary layer thickness with respect to the isolator height decreased as the scale increased. For all the isolator simulations, a shock train was expected to form inside the duct. However, the flow simulation failed to generate this flow pattern, due to improper sizing of the isolator and combustor for a 3-D model or having a low pressure ratio of 2.38. Instead, a single normal shock wave was established at the same relative location within the length of each duct, approximately 80% of the duct length from the isolator entrance. The shape of the shock changed as the scale increased from a normal shock wave, to a bifurcated shock wave, and to a normal shock train, respectively for the 1X, 10X, and 100X models.
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McDaniel, Keith Scott. "Three dimensional simulation of time-dependent scramjet isolator combustor flowfields implemented on parallel architectures /". 2000. http://www.lib.ncsu.edu/etd/public/etd-384420281110043620/etd.pdf.
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Lee, Wei-Shen, i 李維陞. "Simulation and analysis of ethylene-fueled scramjet combustor assisted by a porous cylindrical burner". Thesis, 2018. http://ndltd.ncl.edu.tw/handle/9a5d7m.
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碩士國立臺灣大學
機械工程學研究所
106
This research numerically investigated the flow field in scramjet combustion chamber using licensed free and open source CFD software OpenFOAM. Following future trend of airframe integration design concept, using energy-dense fuel become necessary due to limited fuel storage space, thus hydrocarbon fuel is favored due to higher volumetric energy density, practicality and safety. However, the ignition and flame-holding difficulties, high ignition delay and low flame propagation speed still served as a primary issue. After all, overcoming the drag and producing net thrust. Ethylene has highest volumetric energy density and shortest ignition delay time among gaseous hydrocarbon fuels. Large amount of ethylene will produce during the initial reaction stage of long-chain hydrocarbon liquid fuels (such as kerosene, JP7, JP10 etc.). In order to improve the ignition and flame-holding capability of ethylene-fueled DLR combustor, a cylindrical porous burner was added in downstream of the strut.The drag increase, total pressure loss, flame extinction and flame stabilization mechanism after adding the cylindrical porous burner will be further clarified. Initially, ignition and flame-holding can’t be achieved in original ethylene-fueled DLR combustor. After adding a cylindrical porous burner at L/D = 11.7, ethylene can be ignited and flame-holding can be achieved in cylinder shear layer as a lift-off nonpremixed turbulent flame. High temperature recirculation zone was observed behind the cylinder, which contributes to continuously igniting the mixture in cylinder shear layer. Higher turbulence intensity in shear layer and large amount of small vortex structures in cylinder wake region are observed, which further enhance fuel-air mixing. The porous medium can be heated up to 1000 K and preheating the fuel, which increase the chemical reaction rate and improve combustion stability. However, positive net thrust can’t be produced due to bow shock formation in front of the cylinder which causing too much drag. Further moving the cylinder upward to L/D = 1.7, which located in the wake region of upstream strut. Ethylene can still be ignited and flame can be stabilized in cylinder shear layer. Same flame pattern and flame stabilization mechanism are observed as in L/D = 11.7. In addition, fuel transverse displacement, flame spreading area and oxygen consumption will also increase. Flow shielding effect provided by upstream strut, which prevent shock formation in front of cylinder, thus drag reduction can be achieved. By adding a cylindrical porous burner, ignition of ethylene and flame holding problem can also be solved. Furthermore, higher net thrust and specific impulse can be produced under the same fuel mass flow rate.
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Cassone, Egidio. "Combustion processes modeling: numerical investigations of an engine equipped with a Turbulent Jet Ignition system and of a scramjet combustor". Doctoral thesis, 2022. http://hdl.handle.net/11589/240300.
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Combustion is at the base of many energy production processes and, today like never before, research is involved in maximizing combustion efficiency and reducing the pollution deriving from it. Thanks to the high specific energy of fuels, combustion processes will still have a role in the future, especially in the transportation sector. Therefore, increasing Internal Combustion Engines (ICE) efficiency and reducing their polluting emissions has become an imperative object. Among all the non-conventional ignition systems that are being purposed by research, Turbulent Jet Ignition (TJI) seems to be one of the most promising capable to achieve leaner combustion and higher thermal efficiency in spark ignited engine. In a TJI system a jet of high-energy reactive gases is generated by means of a pilot combustion in a pre-chamber and used to initiate the main combustion event in the cylinder. By virtue of this, TJI devices are able to achieve a more stable combustion also with more problematic fuels, such as Methane.
In the present work, a deep and innovative analysis approach is purposed and applied to a TJI prototype installed on a Methane fueled optically accessible spark ignition research engine. By means of 3D numerical simulations, the behavior of such engine has been monitored and analyzed over a whole engine cycle. Attention has been paid especially on the scavenging, filling and combustion phases in the pre-chamber. Moreover, the jets characteristics and species distribution and evolution are analyzed in order to study the reactive-jet-induced ignition mechanism of the main charge and the associated fuel conversion mechanism. Attention is also given to pollutant species formation and in-cylinder distribution. The purposed approach allowed the characterization of the main phenomena involved in the operation of such system as well as the evaluation of different parameters, such as combustion duration, flame evolution and pre-chamber ignition energy release. The same engine without the pre-chamber has been modeled as well and used for comparison. The bench test data recorded by Istituto Motori di Napoli CNR were employed both to tune the model and to compare the performance increase due to the TJI device.
Parallelly, CFD studies on a scramjet combustor for hypersonic flight purposes fueled with Hydrogen were also carried out. After having validated the model against a set of in-flight representative data, the main physical mechanisms involved in the mixing and combustion processes were analyzed. Moreover, various information about the complex flow patterns and structure were retrieved.
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Singh, Vishal. "Spray Interaction with Supersonic Crossflow". Thesis, 2021. https://etd.iisc.ac.in/handle/2005/5438.
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The flow residence time in a scramjet combustor is of the order of a millisecond (10−3s).
High energy density liquid fuels are the energy carriers of choice in scramjet engines, however liquid fuels must be atomized, evaporated and mixed before heat release by combustion
can occur. Atomization, mixing and ignition require fnite time. Therefore it is important to
study spray formation and its atomization in supersonic flows. Experiments are performed to
study the spray formation from a water jet injected through a plain orifce atomizer with an
exit diameter (d) of 1 mm, into a Mach 2.2 supersonic crossflow. High-speed shadowgraphy
is performed using high-frequency nano-pulsed LASER as well as LED, to capture the
local structures and spray features. The high-speed camera and nano-pulsed LASER is
synchronized using a delay pulse generator. The pulse width of the LASER is kept at 8 and
14 ns, such that to freeze the spray features in time. The bow shock profles are observed
to overlap between different momentum flux ratios (q) of the injected jet when shifted to
the sonic point. The penetration height of the spray is evaluated using the upper spray
boundary. The average spray trajectories are compared for three different momentum flux
ratios. Wavelike disturbances are observed on the windward side of spray, which further
develops into ligaments. Fundamental questions like ligament origin, speed and their breakup
are addressed. The ligaments are present at spray boundary and observed to move with
free stream flow near the injector. Further ligaments are tracked in successive shadowgraph
images using cross-correlation technique to fnd their speed. The ligament speed is also
compared for different momentum flux ratio and found to vary inversely with the momentum flux ratio. The wavelength associated with ligaments or surface waves is observed to increase
linearly along the spray boundary irrespective of the momentum flux ratio. The ligament
breakup is characterized using shear Weber number and wavelength is used as characteristic
length scale. It is found that ligaments break due to shear from free stream flow. Small
shocks (shocklets) are formed ahead of ligaments which are noted to increase the residence
time. Shocklets also delay atomization of the ligaments.
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O'Byrne, Sean Brendan. "Examination of transient mixing and combustion processes in a supersonic combustion ramjet engine". Master's thesis, 1997. http://hdl.handle.net/1885/145993.
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Liu, Wen Jay, i 劉文傑. "Ignition and Transient Combustion Study of Scramjet". Thesis, 2000. http://ndltd.ncl.edu.tw/handle/65168733382412625267.
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博士國立成功大學
航空太空工程學系
88
Ignition and Transient Combustion Study of Scramjet Student : Wen-Jay Liu Advisor : Jir-Ming Char Co-Advisor : Huei-Huang Chiu ABSTRACT Evident technological , economical , and operational advantages exist when liquid hydrocarbon fuels , such as JP-8, RJ-5 are used in comparison with hydrogen-based system for the development of the small hypersonic vehicles. Throughout the continuing development of the scramjet concept, the mixing of fuel and air in the combustor has been one of many important and persistent problems. The length of the combustor must be limited to a few feet due to the impact that the size of the combustor can have on the overall performance of a highly integrated hypersonic flight vehicles. Therefore, the sufficient entraiment and subsquent micromixing followed by significant heat release must all occur in the short residence time of a fuel-air mixture, which is 4), wedges will shorten the ignition delay time to 50 % as compared with no wedges case. The mechanisms of the test section flowfield are pictured and analyzed. Furthermore, the difference between varied combustor configurations will be discussed in detail in this study.
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Star, Jason B. "Numerical simulation of scramjet combustion in a shock tunnel". 2005. http://www.lib.ncsu.edu/theses/available/etd-12082005-152349/unrestricted/etd.pdf.
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Koo, Heeseok. "Large-eddy simulations of scramjet engines". Thesis, 2011. http://hdl.handle.net/2152/ETD-UT-2011-05-3203.
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The main objective of this dissertation is to develop large-eddy simulation (LES) based computational tools for supersonic inlet and combustor design.
In the recent past, LES methodology has emerged as a viable tool for modeling turbulent combustion. LES computes the large scale mixing process accurately, thereby providing a better starting point for small-scale models that describe the combustion process. In fact, combustion models developed in the context of Reynolds-averaged Navier Stokes (RANS) equations exhibit better predictive capability when used in the LES framework. The development of a predictive computational tool based on LES will provide a significant boost to the design of scramjet engines.
Although LES has been used widely in the simulation of subsonic turbulent flows, its application to high-speed flows has been hampered by a variety of modeling and numerical issues. In this work, we develop a comprehensive LES methodology for supersonic flows, focusing on the simulation of scramjet engine components. This work is divided into three sections. First, a robust compressible flow solver for a generalized high-speed flow configuration is developed. By using carefully designed numerical schemes, dissipative errors associated with discretization methods for high-speed flows are minimized. Multiblock and immersed boundary method are used to handle scramjet-specific geometries. Second, a new combustion model for compressible reactive flows is developed. Subsonic combustion models are not directly applicable in high-speed flows due to the coupling between the energy and velocity fields. Here, a probability density function (PDF) approach is developed for high-speed combustion. This method requires solution to a high dimensional PDF transport equation, which is achieved through a novel direct quadrature method of moments (DQMOM). The combustion model is validated using experiments on supersonic reacting flows. Finally, the LES methodology is used to study the inlet-isolator component of a dual-mode scramjet. The isolator is a critical component that maintains the compression shock structures required for stable combustor operation in ramjet mode. We simulate unsteady dynamics inside an experimental isolator, including the propagation of an unstart event that leads to loss of compression. Using a suite of simulations, the sensitivity of the results to LES models and numerical implementation is studied.text
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Karl, Sebastian. "Numerical Investigation of a Generic Scramjet Configuration". Doctoral thesis, 2010. https://tud.qucosa.de/id/qucosa%3A25578.
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A Supersonic Combustion Ramjet (scramjet) is, at least in theory, an efficient air-breathing propulsion system for sustained hypersonic flight at Mach numbers above approximately M=5. Important design issues for such hypersonic propulsion systems, are the lack of ground based facilities capable of testing a full-sized engine at cruise flight conditions and the absence of general scaling laws for the extrapolation of wind tunnel data to flight configurations. Therefore, there is a strong need for the development and validation of CFD tools to support the design process of scramjet-powered vehicles. The aims of this thesis are, in this context, to assess the applicability of, to further develop, and to validate the DLR TAU flow solver for the CFD analysis of the complete flow-path of a scramjet vehicle. The basis of this validation and of the identification of critical modelling assumptions is the recalculation of a series of wind tunnel tests of the HyShot II generic scramjet configuration that were performed in the High Enthalpy Shock Tunnel Göttingen (HEG) at the German Aerospace Center, DLR.Staustrahlantriebe, bei denen sich die Strömung im gesamten Triebwerksbereich im Überschall befindet (supersonic combustion ramjets, Scramjets), stellen ein - zumindest theoretisch - effektives Antriebessystem für den Hyperschallflug im Machzahlbereich von M > 5 dar. Die Auslegung und der Entwurf von luftatmenden Hyperschallantrieben sind in der Praxis mit Schwierigkeiten verbunden. Der Einsatz von Bodenversuchsanlagen ist auf kleinskalige Konfigurationen oder einzelne Triebwerkskomponenten begrenzt. Die Ergebnisse von numerischen Strömungssimulationsverfahren sind mit hohen Unsicherheiten behaftet, die ihren Ursprung in der Modellbildung für die komplexen Strömungsphänomene in chemisch reagierenden, kompressiblen und turbulenten Über- und Hyperschallströmungen haben. Weiterhin existieren keine allgemein gültigen Skalierungsgesetze um Aussagen aus Windkanalexperimenten auf Flugkonfigurationen zu übertragen.Die vorliegende Arbeit beschäftigt sich in diesem Zusammenhang mit der Erweiterung des DLRStrömungslösers TAU für die Berechnung von Überschallverbrennungsphänomenen in Scramjets sowie mit der Anwendung des Verfahrens für die numerische Analyse von Windkanalexperimenten, die im Hochenthalpiekanal Göttingen (HEG) des Deutschen Zentrums für Luft- und Raumfahrt (DLR) zur Untersuchung der generischen HyShot II Scramjet-Konfiguration durchgeführt wurden. Die wichtigsten Ziele waren die genaue Charakterisierung der freien Anströmung im Windkanal, der Nachweis der Anwendbarkeit des verwendeten Rechenverfahrens und die Analyse des Einflusses verschiedener numerischer Modellierungsansätze für die Strömungssimulation in Scramjets sowie die Nutzung der numerischen Daten für eine verbesserte Interpretation der experimentellen Ergebnisse.
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Cheng, Yu-Chun, i 鄭伃均. "Simulation and analysis of scramjet combustion chamber assisted by a porous cylindrical burner". Thesis, 2016. http://ndltd.ncl.edu.tw/handle/71763092032844728750.
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碩士國立臺灣大學
機械工程學研究所
104
The study is based on the theory of gas dynamics. We use the software FLUENT to simulate the flow field scramjet combustion chamber. We choosed SST k-ωas turbulence model, because it can get solution which is more close to real flow field. The simulation procedure is verified against by two examples to ensure its accuracy, and we move on to design scramjet combustion chamber. Adding a porous cylinder as a burner into the flow field to improve its performance,including flame stabilization, kinetic energy and thrust improvement. Using different inlet condition to compare with the original flow field. We take some limitations into account, including the drag effect to the flow field, and flame extinction. The pressure at combustion chamber outlet should be higher than nozzle outlet for the purpose of continuous flow. The temperature at combustion chamber should avoid hot spot or it may cause damage on it. Considering the above conditions, we ensure that the performance of combustion chamber assisted by porous cylinder do better than the original one by checking the steady simulation. In the end, it could be shown that the combustion chamber assisted by porous cylinder could be ignited at extremely low hydrogen mass flow rate, which cannot be done on the original one.