Dissertations / Theses on the topic 'Supersonic combustion'
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1
Lou, Zhipeng. "Improved Flamelet Modeling of Supersonic Combustion." Thesis, State University of New York at Stony Brook, 2017. http://pqdtopen.proquest.com/#viewpdf?dispub=10280296.
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A computational fluid dynamics (CFD)-based study using large-eddy simulation (LES) and the flamelet-progress variable (FPV) approach for turbulence-combustion interaction has been undertaken to investigate the combustion that takes place under supersonic flow conditions. The target application is the propulsive system associated with dual-mode scramjet, which has been recognized as the most promising air-breathing system for hypersonic flight. In addition to the standard practice of using mixture fraction and its dissipation rate as independent variables of the look-up table in the flamelet procedure for non-premixed flames, pressure has been added to enable the inclusion of its effects on chemical reactions under high speed conditions. An improved method of generating the flamelet library that allows new interpolations based on the three branches of the reaction curve (S-Curve) in non-premixed combustion has been proposed during the course of the present work. Solutions of supersonic combustion in three different configurations have been used to assess the accuracy of the various proposed improvements and investigate fundamental physics of dual-mode scramjets.
2
Luo, Wenlei. "Large Eddy Simulation of turbulent supersonic combustion and characteristics of supersonic flames." Thesis, University of Leeds, 2014. http://etheses.whiterose.ac.uk/7641/.
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In this thesis we investigate the supersonic combustion in scramjet combustors with strut and cavity flame holders through the Reynolds-Averaged Navier–Stokes (RANS) and Large Eddy Simulation (LES) strategies. Firstly, the Unsteady Flamelet/Progress Variable (UFPV) model for turbulent combustion in low-speed flows is introduced and extended to supersonic flows and a new strategy is developed to create probability density function look-up tables for the UFPV model. Secondly, the RANS modelling is employed to a strut-based scramjet combustor using the flamelet and UFPV models and the latter shows a better performance. Subsequently, the LES modelling is performed with the UFPV model and the UFPV model gives good predictions on comparing the numerical results to the experimental data. Thirdly, the LES modelling is employed to a cavity-based scramjet combustor. The results obtained indicate that the local extinction and autoignition events are very common phenomena in the supersonic flame and the UFPV model is able of predicting these events with reasonable accuracy. Further, an activation-energy-asymptotic-based Damköhler number concept is a valuable metric to identify flame weakening and extinction in supersonic flames. Together with the OH radicals, the distribution of the HO2 radicals can assist in identifying the autoignition events in the supersonic flame. Finally, analysing the flameholding mechanisms of the cavity, it is found that the cavity provides a stable ignition source to the fluid. Further, the combustion in the cavity is dominated by flame propagation. However, on the outer interface of the air and hydrogen streams, the combustion is mainly dominated by autoignition. Both autoignition and flame propagation contribute to the combustion in the mixing layer. Also the combustion in the cavity mixing layer has effects on the induction reactions in the wake of the hydrogen jet and reduces the induction time of the autoignition.
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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.
4
Do, Hyungrok. "Plasma-assisted combustion in a supersonic flow /." May be available electronically:, 2009. http://proquest.umi.com/login?COPT=REJTPTU1MTUmSU5UPTAmVkVSPTI=&clientId=12498.
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Picciani, Mark. "Supersonic combustion modelling using the conditional moment closure approach." Thesis, Cranfield University, 2014. http://dspace.lib.cranfield.ac.uk/handle/1826/9309.
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This work presents a novel algorithm for supersonic combustion modelling. The method involved
coupling the Conditional Moment Closure (CMC) model to a fully compressible, shock
capturing, high-order flow solver, with the intent of modelling a reacting hydrogen-air, supersonic
jet.
Firstly, a frozen chemistry case was analysed to validate the implementation of the algorithm
and the ability for CMC to operate at its frozen limit. Accurate capturing of mixing is crucial
as the mixing and combustion time scales for supersonic flows are on the order of milliseconds.
The results of this simulation were promising even with an unexplainable excess velocity decay
of the jet core. Hydrogen mass fractions however, showed fair agreement to the experiment.
The method was then applied to the supersonic reacting case of ONERA. The results showed
the method was able to successfully capture chemical non-equilibrium effects, as the lift-off
height and autoignition time were reasonably captured. Distributions of reactive scalars were
difficult to asses as experimental data was deemed to be very inaccurate. As a consequence,
published numerical results for the same test case were utilised to aid in analysing the results of
the presented simulations. Due to the primary focus of the study being to assess non-equilibrium
effects, the clustering of the computational grid lent itself to smeared and lower magnitude wall
pressure distributions. Nevertheless, the wall pressure distributions showed good qualitative
agreement to experiment.
The primary conclusions from the study were that the CMC method is feasible to model
supersonic combustion. However, a more detailed analysis of sub-models and closure assumptions
must be conducted to assess the feasibility on a more fundamental level. Also, from the
results of both the frozen chemistry and the reacting case, the effects of assuming constant
species Lewis number was visible.
6
Makowka, Konrad [Verfasser]. "Numerically Efficient Hybrid RANS/LES of Supersonic Combustion / Konrad Makowka." München : Verlag Dr. Hut, 2016. http://d-nb.info/1084385236/34.
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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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Tedder, Sarah Augusta. "Advancements in dual-pump broadband CARS for supersonic combustion measurements." W&M ScholarWorks, 2010. https://scholarworks.wm.edu/etd/1539623572.
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Space- and time-resolved measurements of temperature and species mole fractions of nitrogen, oxygen, and hydrogen were obtained with a dual-pump coherent anti-Stokes Raman spectroscopy (CARS) system in hydrogen-fueled supersonic combustion free jet flows. These measurements were taken to provide time-resolved fluid properties of turbulent supersonic combustion for use in the creation and verification of computational fluid dynamic (CFD) models. CFD models of turbulent supersonic combustion flow currently facilitate the design of air- breathing supersonic combustion ramjet (scramjet) engines. Measurements were made in supersonic axi-symmetric free jets of two scales. First, the measurement system was tested in a laboratory environment using a laboratory-scale burner (∼10 mm at nozzle exit). The flow structures of the laboratory-burner were too small to be resolved with the CARS measurements volume, but the composition and temperature of the jet allowed the performance of the system to be evaluated. Subsequently, the system was tested in a burner that was approximately 6 times larger, whose length scales are better resolved by the CARS measurement volume. During both these measurements, weaknesses of the CARS system, such as sensitivity to vibrations and beam steering and inability to measure temperature or species concentrations in hydrogen fuel injection regions were identified. Solutions were then implemented in improved CARS systems. One of these improved systems is a dual-pump broadband CARS technique called, Width Increased Dual-pump Enhanced CARS (WIDECARS). The two lowest rotational energy levels of hydrogen detectable by WIDECARS are H2 S(3) and H2 S(4). The detection of these lines gives the system the capability to measure temperature and species concentrations in regions of the flow containing pure hydrogen fuel at room temperature. WIDECARS is also designed for measurements of all the major species (except water) in supersonic combustion flows fueled with hydrogen and hydrogen/ethylene mixtures (N2, O 2, H2, C2H4, CO, and CO2). This instrument can characterize supersonic combustion fueled with surrogate fuel mixtures of hydrogen and ethylene. This information can lead to a better understanding of the chemistry and performance of supersonic combustion fueled with cracked jet propulsion (JP)-type fuel.
9
Sexton, Scott Michael. "Progress Toward Analytic Predictions of Supersonic Hydrocarbon-Air Combustion| Computation of Ignition Times and Supersonic Mixing Layers." Thesis, University of California, San Diego, 2018. http://pqdtopen.proquest.com/#viewpdf?dispub=10687717.
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Combustion in scramjet engines is faced with the limitation of brief residence time in the combustion chamber, requiring fuel and preheated air streams to mix and ignite in a matter of milliseconds. Accurate predictions of autoignition times are needed to design reliable supersonic combustion chambers. Most efforts in estimating non-premixed autoignition times have been devoted to hydrogen-air mixtures. The present work addresses hydrocarbon-air combustion, which is of interest for future scramjet engines.
Computation of ignition in supersonic flows requires adequate characterization of ignition chemistry and description of the flow, both of which are derived in this work. In particular, we have shown that activation energy asymptotics combined with a previously derived reduced chemical kinetic mechanism provides analytic predictions of autoignition times in homogeneous systems. Results are compared with data from shock tube experiments, and previous expressions which employ a fuel depletion criterion.
Ignition in scramjet engines has a strong dependence on temperature, which is found by perturbing the chemically frozen mixing layer solution. The frozen solution is obtained here, accounting for effects of viscous dissipation between the fuel and air streams. We investigate variations of thermodynamic and transport properties, and compare these to simplified mixing layers which neglect these variations. Numerically integrating the mixing layer problem reveals a nonmonotonic temperature profile, with a peak occurring inside the shear layer for sufficiently high Mach numbers.
These results will be essential in computation of ignition distances in supersonic combustion chambers.
10
Billingsley, Matthew C. "Plasma Torch Atomizer-Igniter for Supersonic Combustion of Liquid Hydrocarbon Fuels." Thesis, Virginia Tech, 2005. http://hdl.handle.net/10919/36331.
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To realize supersonic combustion of hydrocarbons, an effective atomizer-igniter combination with the capabilities of fuel preheating, atomization, penetration, mixing, ignition and flameholding is desired. An original design concept incorporating these capabilities was built and tested at Virginia Tech, and was found to provide good penetration, effective atomization, and robust ignition and flameholding. Quiescent testing with kerosene and JP-7 provided initial performance data. The atomizer-injector design was then modified for insertion into a supersonic wind tunnel, and tested with kerosene in an unheated Mach 2.4 flow with typical freestream conditions of To = 280 K and Po = 360 kPa. Water injection was utilized in both cases for comparison and to analyze atomization behavior.
In the quiescent environment, the regeneratively cooled plasma torch igniter was found to significantly increase electrode life while heating, atomizing, and igniting the liquid fuel. Jet breakup length was measured and characterized, and mean droplet size was estimated using an existing correlation. Several qualitative observations regarding quiescent combustion were made, including torch power effects and the process of flame formation. In the supersonic environment, the effect of fuel injection direction was analyzed. Best results were obtained when fuel was injected with a velocity component opposite to the direction of main tunnel flow. Repeatable ignition occurred in the supersonic boundary layer at the fuel stagnation location near the plasma torch plume. Direct, filtered, shadowgraph, and schlieren photographs, temperature measurements, and visible emission spectroscopy provided evidence of combustion and the details of the flame structure.
The new atomizer-igniter design provided robust and reliable ignition and flameholding of liquid hydrocarbon fuels in an unheated supersonic flow at M=2.4, with no ramp, step, or other physical penetration into the flowpath.Master of Science
11
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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Barbi, Eric. "Uncooled choked plasma torch for ignition and flameholding in supersonic combustion." Thesis, Virginia Tech, 1986. http://hdl.handle.net/10919/45744.
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An experimental investigation on an uncooled choked plasma torch using hydrogen/argon mixtures as a propellant was conducted. This low-power plasma torch was designed to be used as an ignitor and flameholder in supersonic combustion. The anode and cathode were made of two-percent thoriated tungsten, and no cooling system was required. Sonic flow through the nozzle was obtained by using a small throat diameter (0.813 mm).
The plasma torch can operate stably over a wide range hydrogen/argon mixtures at power levels of 500 to 2000 W. Voltage-current characteristics of the arc are presented for discharge currents ranging from 5 to 40 A and for various flow rates and mixture fractions. The electrical input power is found to be a linear function of the hydrogen flow for a constant argon flow and for a current of 20 A. Measurements with a calorimeter reveal that the thermal efficiency, defined as the rate of increase of total enthalpy of the gas flowing through the torch divided by the electrical input power, is about 88 percent.Master of Science
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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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Mozingo, Joseph Alexander. "Evaluation of a Strut-Plasma Torch Combination as a Supersonic Igniter-Flameholder." Thesis, Virginia Tech, 2004. http://hdl.handle.net/10919/36461.
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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
15
Gessner, Thomas. "Dynamic mesh adaption for supersonic combustion waves modeled with detailed reaction mechanisms." [S.l. : s.n.], 2001. http://www.freidok.uni-freiburg.de/volltexte/292.
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DeTurris, Dianne Joan. "A technique for direct measurement of skin friction in supersonic combustion flow." Diss., Virginia Tech, 1992. http://hdl.handle.net/10919/39449.
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Blot, Adrian. "Design of a non-vitiated heater ground test facility for supersonic combustion." [Gainesville, Fla.] : University of Florida, 2009. http://purl.fcla.edu/fcla/etd/UFE0024145.
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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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Cross, Melissa A. "Operation Characteristics of a Plasma Torch for Supersonic Combustion Applications with Simulated Cracked JP-7 Feedstock." Thesis, Virginia Tech, 2004. http://hdl.handle.net/10919/42875.
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Research conducted at Virginia Tech has examined plasma torch operational characteristics using a feedstock gas of mixed hydrocarbons representing a cracked JP-7 surrogate. The tests were part of a program to examine the torch as an igniter and flame-holder for hydrocarbon-fueled scramjet engines. Previous research has shown that the plasma torch has promise as a robust igniter and flame-holder using gaseous fuels such as methane, ethylene and propylene when combined with an aeroramp to assist with the combustion process. The present investigation tested the plasma torch with a feedstock mixture of gaseous hydrocarbons that simulates thermally cracked JP-7 jet fuel. This simulation of a cracked hydrocarbon fuel was studied to lay the foundation for work with liquid hydrocarbon fuel, which is of interest for todayâ s aerospace vehicles. The cracked JP-7 surrogate consists of a 15/25/60 mixture of methane/ethane/ethylene. The research results include torch operational characteristics such as downstream plume temperatures and emission spectroscopy within the combustion plume, as well as the power supplied to the torch over a range of mass flow rates. Filtered photographs of the emissions plume were studied to aid torch plume diagnostics. Other observations made were erosion and alignment of the electrodes, which will help determine the potential lifespan of the torch using cracked JP-7 fuel. The results show successful operation over a range of powers with simulated cracked JP-7 feedstock flows. Measured spectra, current, and voltage are compared with similar results for other hydrocarbon feedstock gases. The torch operating on the JP-7 surrogate feedstock appears to be a satisfactory device for ignition, flame-holding, and combustion enhancement of cracked hydrocarbons in supersonic combustion.Master of Science
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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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Gallimore, Scott Douglas. "A Study of Plasma Ignition Enhancement for Aeroramp Injectors in Supersonic Combustion Applications." Diss., Virginia Tech, 2001. http://hdl.handle.net/10919/26988.
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The main goal of this project was to investigate the mixing and chemical phenomena associated with the integration of a low-power, uncooled plasma torch into a fuel injector array. The potential application was for an integrated scramjet igniter/injector, with the hope of producing superior mixing and flameholding performance for supersonic combustion applications. To create a knowledge base for integration, several key investigations were made of the anode material, anode geometry, and spectrographic analysis of different light hydrocarbon fuels and inert feedstocks, all aimed at increasing the ignition potential of the plasma torch. Investigations of the anode material demonstrated the molybdenum provided longer lifetimes than either pure copper or tungsten-copper anodes. In addition, geometric studies of the anode revealed that anodes with short constrictor lengths and sonic exit nozzles provided superior ignition performance based on higher transfer rates of thermal energy from the arc to the feedstock. This resulted in the production of higher hydrogen atom concentrations within the plasma jet. Spectrographic observation of the plasma jets revealed that methane, ethylene, propylene, and propane plasmas all contain excited atomic hydrogen, a radical known to participate in important chain-branching combustion reactions.
Based on the knowledge gained, and encouraging results, a candidate scramjet igniter and flameholder was designed. The design was observed to exhibit a synergistic effect between the plasma igniter and fuel injector in that the fuel injector provides not only a subsonic region for plasma ignition, but also lifts the combustion enhancing radicals out into the fuel-air stream by means of counter-rotating vortices. Furthermore, under the conditions tested, increases in plasma torch power produced an exponential increase in the intensity of downstream products, indicating an enhancement effect. Based upon these observations, the integrated igniter/injector design is expected to perform well in supersonic combustion applications.Ph. D.
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Genin, Franklin Marie. "Study of compressible turbulent flows in supersonic environment by large-eddy simulation." Diss., Atlanta, Ga. : Georgia Institute of Technology, 2009. http://hdl.handle.net/1853/28085.
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Thesis (M. S.)--Aerospace Engineering, Georgia Institute of Technology, 2009.Committee Chair: Menon, Suresh; Committee Member: Ruffin, Stephen; Committee Member: Sankar, Lakshmi; Committee Member: Seitzman, Jerry; Committee Member: Stoesser, Thorsten
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Stouffer, Scott David. "The development and operating characteristics of an improved plasma torch for supersonic combustion applications." Thesis, Virginia Polytechnic Institute and State University, 1989. http://hdl.handle.net/10919/76046.
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The design of the VPI plasma torch, which has been used as an ignitor and flameholder in supersonic combustion studies, has been modified in order to decrease the electrode wear and to increase stability. The plasma torch can be used as a source of hydrogen or nitrogen radicals which initiate and stabilize combustion. During previous testing of the unmodified torch, electrode erosion limited operation of the torch to about two hours.
The improved torch features a flow swirler in the gas inlet, which adds vortex stabilization to the arc. The vortex stabilization causes the anode attachment point of the arc to be anchored in the low pressure region, downstream of the constrictor. This lowers the heat flux to the anode, so that erosion is decreased. The torch body was redesigned with an emphasis on the alignment of the electrodes. Also, the electrode gap in the improved torch was made continuously adjustable, allowing fine adjustment of the electrode gap during operation of the torch.
The operational characteristics of the improved torch were monitored by a microcomputer-based data acquisition system. Stable operation of the improved torch with pure nitrogen was demonstrated, thus eliminating the requirement for argon to stabilize the arc. Operational characteristics of the improved torch running on argon, nitrogen, argon/hydrogen and argon/nitrogen mixtures as feedstocks, are reported. The electrode wear was studied between tests by observation with a microscope, and by measuring the mass change of the electrodes. The electrode erosion of the improved torch was reduced significantly. Anode lifetimes of greater than 20 hours have been demonstrated with operation on mixtures of nitrogen and argon.Master of Science
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Wagner, Timothy Charles. "Ignition and flameholding in supersonic flow by injection of dissociated hydrogen." Diss., Virginia Polytechnic Institute and State University, 1987. http://hdl.handle.net/10919/49905.
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The objective of this research was to investigate analytically and experimentally the use of free radicals for ignition and flameholding in supersonic flows. An analytical investigation of the effects of adding small quantities of radicals to a stoichiometric mixture of hydrogen and air was performed using a finite-rate chemical kinetics code. The results of these calculations indicate that small additions of hydrogen atoms, oxygen atoms, nitrogen atoms, or hydroxyl radicals are effective in promoting ignition. These analytical results were qualitatively verified in a Mach 2 flow experiment using hydrogen atoms generated by a plasma torch. The supersonic combustion tests were conducted in a direct-connect mode at atmospheric pressure with either ambient temperature air or burner-heated vitiated air with total temperatures from 1200 to 4000 R. Both semi-freejet and ducted configurations were used. The experimental results indicate that hydrogen atoms from a low-power plasma torch provide an effective ignition and flameholding source for hydrogen-fueled Mach 2 flows at total temperatures as low as 1065 R, the lowest temperature tested. A reduction in the minimum total temperature required for ignition of several hydrocarbon fuels was also demonstrated.
A piloted fuel injector configuration designed to take maximum advantage of the hydrogen atoms from the plasma torch was conceived and fabricated. The injector design consisted of five small upstream pilot fuel injectors, a rearward-facing step and three primary fuel injectors downstream of the step. The hydrogen atoms from the plasma torch were injected in the recirculation region downstream of the step. Three other ignition sources were also tested as comparisons: an argon plasma, a pyrophoric mixture of silane and hydrogen, and a surface discharge device. Hydrogen-fueled supersonic combustion tests were conducted at conditions similar to those described earlier. Hydrogen atoms generated by the plasma torch proved to be the most effective ignition source, causing ignition for a torch input power of 780 W, the lowest power tested. The combination of the hydrogen atoms and the piloted fuel injector was shown to be a very effective igniter and flameholder for scramjet operation over a simulated flight envelope (Mach 3 to Mach 6, low to moderate altitudes).Ph. D.
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Jacobsen, Lance Steven. "An Integrated Aerodynamic-Ramp-Injector/ Plasma-Torch-Igniter for Supersonic Combustion Applications with Hydrocarbon Fuels." Diss., Virginia Tech, 2001. http://hdl.handle.net/10919/28858.
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The first integrated, flush-wall, aero-ramp-fuel-injector/plasma-torch igniter and flame propagation system for supersonic combustion applications with hydrocarbon fuels was developed and tested. The main goal of this project was to develop a device which could be used to demonstrate that the correct placement of a plasma-torch-igniter/flame-holder in the wake of the fuel jets of an aero-ramp injector array could make sustained, efficient supersonic combustion with low losses and thermal loading possible in a high enthalpy environment. Three phases of research were performed to develop the device using the supersonic cold flow facilities at Virginia Tech. The experimental investigations included some of the following methods: shadowgraphs, surface oil flow, pressure-sensitive paint, high- or low-speed photography, aerothermodynamic sampling, and spectroscopy. During this research effort, a new mixing parameter was also developed to quantify the injector plume mass fraction concentration values using successive profiles of ambient or heated air as the injectant.
The first phase of the research effort was conducted at Mach 3.0 at a static pressure and temperature of 0.19 atm and 101 K. This phase involved component analyses to improve on the designs of the aero-ramp and plasma-torch as well as address integration and incorporation difficulties. The information learned from these experiments lead to the creation of the first prototype integrated aero-ramp/plasma torch design featuring a new simplified four-hole aero-ramp design.
The second phase of the project consisted of experiments at Mach 2.4 involving a cold-flow mixing evaluation of the new aero-ramp design and a resizing of the device for incorporation into a scramjet flow path test rig at the Air Force Research Laboratories (AFRL). Experiments were performed at a static pressure and temperature of 0.25 atm and 131 K, and at injector-jet to freestream momentum flux ratios ranging from 1.0 to 3.3. Results showed the aero-ramp to mix at a considerably faster rate than the injector used in the AFRL baseline combustor configuration due to high levels of vorticity created by the injector array. In addition, the plume of the aero-ramp lifted off the test section wall without trapping a secondary core inside the shear layer near the surface, unlike the earlier nine-hole aero-ramp arrays. The mitigation of the secondary fuel core leads to a lower level of combustion near the surface and a lower potential for thermal loading on the wall.
The last phase of the research involved testing the final device design in a cold-flow environment at Mach 2.4 with ethylene fuel injection and an operational plasma torch with methane, nitrogen, a 90-percent nitrogen 10-percent hydrogen (by volume) mixture, and air feedstock gases. Experiments were performed with injector jet to freestream momentum flux ratios ranging from 1.4 to 3.3, and 1.2 with the plasma torch at a nominal power level 2000 watts. Overall, the final integrated design showed a high mixing efficiency and a higher potential for repeatable main fuel ignition and flame propagation with the plasma torch placed at the middle of the three downstream torch stations tested (x/dinjector = 8 downstream from the center of the injector area), with nitrogen as the torch feedstock. Furthermore, the integrated device created a sustained flame, demonstrating main fuel ignition in a cold and low pressure supersonic environment with a plasma-torch. Local intensity distributions of the major excited species generated from the interaction of the plasma-torch with the main fuel plume were also identified with a spectrometer. As a result of the research and development process, an injector block for scramjet combustor experiments consisting of four integrated aero-ramp-injector/plasma-torch-igniters was created for near future tests at the AFRL.Ph. D.
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Wickersham, Andrew Joseph. "Development of Multi-perspective Diagnostics and Analysis Algorithms with Applications to Subsonic and Supersonic Combustors." Diss., Virginia Tech, 2014. http://hdl.handle.net/10919/51145.
Full textAbstract:
There are two critical research needs for the study of hydrocarbon combustion in high speed flows: 1) combustion diagnostics with adequate temporal and spatial resolution, and 2) mathematical techniques that can extract key information from large datasets. The goal of this work is to address these needs, respectively, by the use of high speed and multi-perspective chemiluminescence and advanced mathematical algorithms.
To obtain the measurements, this work explored the application of high speed chemiluminescence diagnostics and the use of fiber-based endoscopes (FBEs) for non-intrusive and multi-perspective chemiluminescence imaging up to 20 kHz. Non-intrusive and full-field imaging measurements provide a wealth of information for model validation and design optimization of propulsion systems. However, it is challenging to obtain such measurements due to various implementation difficulties such as optical access, thermal management, and equipment cost. This work therefore explores the application of FBEs for non-intrusive imaging to supersonic propulsion systems. The FBEs used in this work are demonstrated to overcome many of the aforementioned difficulties and provided datasets from multiple angular positions up to 20 kHz in a supersonic combustor. The combustor operated on ethylene fuel at Mach 2 with an inlet stagnation temperature and pressure of approximately 640 degrees Fahrenheit and 70 psia, respectively. The imaging measurements were obtained from eight perspectives simultaneously, providing full-field datasets under such flow conditions for the first time, allowing the possibility of inferring multi-dimensional measurements.
Due to the high speed and multi-perspective nature, such new diagnostic capability generates a large volume of data and calls for analysis algorithms that can process the data and extract key physics effectively. To extract the key combustion dynamics from the measurements, three mathematical methods were investigated in this work: Fourier analysis, proper orthogonal decomposition (POD), and wavelet analysis (WA). These algorithms were first demonstrated and tested on imaging measurements obtained from one perspective in a sub-sonic combustor (up to Mach 0.2). The results show that these algorithms are effective in extracting the key physics from large datasets, including the characteristic frequencies of flow—flame interactions especially during transient processes such as lean blow off and ignition. After these relatively simple tests and demonstrations, these algorithms were applied to process the measurements obtained from multi-perspective in the supersonic combustor. compared to past analyses (which have been limited to data obtained from one perspective only), the availability of data at multiple perspective provide further insights into the flame and flow structures in high speed flows.
In summary, this work shows that high speed chemiluminescence is a simple yet powerful combustion diagnostic. Especially when combined with FBEs and the analyses algorithms described in this work, such diagnostics provide full-field imaging at high repetition rate in challenging flows. Based on such measurements, a wealth of information can be obtained from proper analysis algorithms, including characteristic frequency, dominating flame modes, and even multi-dimensional flame and flow structures.Ph. D.
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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.
Full textAbstract:
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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Vellaramkalayil, Jiby Jacob [Verfasser]. "Experimental and Numerical Investigations of Different Injection Schemes in a Supersonic Combustion Chamber / Jiby Jacob Vellaramkalayil." München : Verlag Dr. Hut, 2014. http://d-nb.info/1058285467/34.
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Rabadán, Santana Edder José [Verfasser]. "Numerical Investigation of a Generic Supersonic Combustion Chamber under Realistic Flight Conditions / Edder José Rabadán Santana." München : Verlag Dr. Hut, 2015. http://d-nb.info/1074063570/34.
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Tabbara, Hani. "Numerical investigations of thermal spray coating processes : combustion, supersonic flow, droplet injection, and substrate impingement phenomena." Thesis, University of Southampton, 2012. https://eprints.soton.ac.uk/348993/.
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The aim of this thesis is to apply CFD methods to investigate the system characteristics of high speed thermal spray coating processes in order facilitate technological development. Supersonic flow phenomena, combustion, discrete droplet and particle migration with heating, phase change and disintegration, and particle impingement phenomena at the substrate are studied. Each published set of results provide an individual understanding of the underlying physics which control different aspects of thermal spray systems. A wide range of parametric studies have been carried out for HVOF, warm spray, and cold spay systems in order to build a better understanding of process design requirements. These parameters include: nozzle cross-section shape, particle size, processing gas type, nozzle throat diameter, and combustion chamber size. Detailed descriptions of the gas phase characteristics through liquid fuelled HVOF, warm spray, and cold spray systems are built and the interrelations between the gas and powder particle phases are discussed. A further study looks in detail at the disintegration of discrete phase water droplets, providing a new insight to the mechanisms which control droplet disintegration, and serves as a fundamental reference for future developments of liquid feedstock devices. In parallel with these gas-particle-droplet simulations, the impingement of molten and semi-molten powder droplets at the substrate is investigated and the models applied simulate the impingement, spreading and solidification. The results obtained shed light on the break-up phenomena on impact and describe in detail how the solidification process varies with an increasing impact velocity. The results obtained also visually describe the freezing induced break-up phenomenon at the splat periphery.
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Guven, Umut. "Simulation haute-fidélité de la combustion pour les moteurs-fusées." Thesis, Normandie, 2018. http://www.theses.fr/2018NORMIR30/document.
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L’allumage est un point essentiel dans le dimensionnement des moteurs-fusées, et il nécessite de prendre en compte plusieurs phénomènes physiques très distincts qui sont autant de challenges numériques. Le premier point abordé pendant cette thèse est la modélisation et la simulation par Simulation aux Grandes Échelles d’un allumeur de type VINCI. Des gaz chauds, riches en oxygène, sont délivrés de façon supersonique dans une chambre remplie d’hydrogène faisant apparaître un jet fortement sous-détendu et de multiples interactions choc/choc ou choc/flamme. Les premiers instants du processus d’allumage sont ici détaillés. Le second point abordé est la modélisation et la simulation numérique de la combustion H2/O2 à haute pression. En particulier, les effets d’une diffusion non-idéale sont étudiés dans le cas de flammes de prémélange 1D et sur la configuration 2D de type ‘splitter plate’. Un impact de la modélisation sur les espèces produites et le champ de température est ici mis en lumièreIgnition is a key point in the design of liquid rocket engine (LRE), and it requires to take into account several distinct physical phenomena that constitute numerical challenges. The first point addressed during this thesis is the modeling and simulation using Large Eddy Simulation of a LRE igniter in a configuration close to VINCI rocket engine. The hot gases from the igniter, rich in oxygen, are delivered at supersonic speeds in a chamber filled with hydrogen. Such configuration creates under-expanded jets with multiple shock/shock or shock/flame interactions. A focus is done on the ignition process. The second point addressed is the modeling and simulation of high pressure H2/O2 combustion which occurs. In particular, the effects of non-ideal diffusion are studied through a 1D premixed flames and a 2D splitter plate configuration. An impact of modeling on the species produced and the temperature field is highlighted
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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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Fortier-Topping, Hugo. "Conception d'une chambre de combustion pour la microturbine à gaz SRGT-2." Mémoire, Université de Sherbrooke, 2014. http://hdl.handle.net/11143/5417.
Full textAbstract:
Dans un contexte mondial où les ressources énergétiques commencent à se faire rares, beaucoup de recherches se font sur l’amélioration de l’efficacité thermique et de la densité de puissance des sources d’énergie existantes. Ainsi, un projet de développement d’une microturbine à gaz avec une architecture de nouveau genre permettant d’augmenter la densité de puissance tout en réduisant les coûts a vu le jour.
La recherche proposée dans le présent document se concentre sur la conception et la caractérisation d’une chambre de combustion et d’un banc d’essai pour la turbine SRGT-2.
Une chambre de combustion à écoulement inverse est conçue et caractérisée expérimentalement. Un modèle 0D de la chambre est tout d’abord fait. Par la suite, une optimisation numérique est faite jusqu’à l’atteinte des objectifs de conception. Finalement, la chambre de combustion est testée durant 30 secondes avec de l’hydrogène comme carburant. Une température de sortie de la chambre de combustion de 1000 K a été maintenue avec une efficacité de combustion de plus de 85%.
Le banc d’essai conçu pour le projet de recherche utilise un démarreur électropneumatique permettant d’accélérer le prototype jusqu’à 102 000 RPM. Le module fluide est la partie du banc d’essai qui contient les différentes parties de la turbine SRGT-2 comme le rotor, les stators et la chambre de combustion. Le module est instrumenté dans le but d’obtenir une caractérisation complète de la turbine. Sa configuration modulaire permet aussi de caractériser chacune des composantes individuellement en changeant certaines sections.
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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.
Full textAbstract:
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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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.
Full textAbstract:
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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Griffiths, Alan David, and 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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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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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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Wittmers, Nicole K. "Direct-connect performance evaluation of a valveless pulse detonation engine." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2004. http://library.nps.navy.mil/uhtbin/hyperion/04Dec%5FWittmers.pdf.
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Martínez, Ferrer Pedro José. "Étude par simulation numérique de l'auto-allumage en écoulement turbulent cisaillé supersonique." Thesis, Chasseneuil-du-Poitou, Ecole nationale supérieure de mécanique et d'aérotechnique, 2013. http://www.theses.fr/2013ESMA0018.
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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 similairesThis 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
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Rocci, Denis Sara [Verfasser], Hans-Peter [Akademischer Betreuer] Kau, Thomas [Akademischer Betreuer] Sattelmayer, and Christian [Akademischer Betreuer] Mundt. "Design and Experimental Study of Injection Systems in a Supersonic Combustion Chamber / Sara Rocci Denis. Gutachter: Thomas Sattelmayer ; Christian Mundt. Betreuer: Hans-Peter Kau." München : Universitätsbibliothek der TU München, 2011. http://d-nb.info/1019588853/34.
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Melen, Stéphane. "Modélisation et étude numérique de la combustion supersonique turbulente non-prémélangée, approche probabiliste." Rouen, 1995. http://www.theses.fr/1995ROUE5044.
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Après avoir présenté le problème physique, la première partie décrit les équations et les modèles utilisés pour simuler un écoulement réactif, turbulent et supersonique. La description de la turbulence fait appel à un modèle k- incluant les effets de compressibilité et les effets dus au dégagement de la chaleur. La modélisation des effets de chimie non-infiniment rapide dans un écoulement turbulent est assurée par une approche fonction densité de probabilité. Le modèle PEUL-diffusion a été retenu et introduit dans le code de calcul. Ce modèle fournit une forme présumée de la PDF des variables thermochimiques. Dans la deuxième partie de ce travail sont présentées les méthodes numériques utilisées pour la partie Navier-Stokes et pour le modèle PEUL. Pour résoudre la partie eulérienne, un schéma explicite/implicite inconditionnellement stable TVD a été choisi. Dans la dernière partie, deux simulations de flamme jet d'hydrogène dans l'air sont comparées avec des résultats expérimentaux
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Poubeau, Adèle. "Simulation des émissions d'un moteur à propergol solide : vers une modélisation multi-échelle de l'impact atmosphérique des lanceurs." Thesis, Toulouse 3, 2015. http://www.theses.fr/2015TOU30039/document.
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Les lanceurs ont un impact sur la composition de l'atmosphere, et en particulier sur l'ozone stratospherique. Parmi tous les types de propulsion, les moteurs à propergol solide ont fait l'objet d'une attention particulière car leurs émissions sont responsables d'un appauvrissement significatif d'ozone dans le panache des lanceurs lors des premières heures suivant le lancement. Ce phénomène est principalement dû à la conversion de l'acide chlorhydrique, un composé chimique présent en grandes quantités dans les émissions de ce type de moteur, en chlore actif qui réagit par la suite avec l'ozone dans un cycle catalytique similaire à celui responsable du "trou de la couche d'ozone", cette diminution périodique de l'ozone en Antarctique. Cette conversion se produit dans le panache supersonique, où les hautes températures favorisent une seconde combustion entre certaines espèces chimiques du panache et l'air ambiant. L'objectif de cette étude est d'évaluer la concentration de chlore actif dans le panache d'un moteur à propergol solide en utilisant la technique des Simulations aux Grandes Echelles (SGE). Le gaz est injecté à travers la tuyère d'un moteur et une méthode de couplage entre deux instances du solveur de mécanique des fluides est utilisée pour étendre autant que possible le domaine de calcul derrière la tuyère (jusqu'à l'équivalent de 400 diamètres de sortie de la tuyère). Cette méthodologie est validée par une première SGE sans chimie, en analysant les caractéristiques de l'écoulement supersonique avec co-écoulement obtenu par ce calcul. Ensuite, le chimie mettant en jeu la conversion des espèces chlorées a été étudiée au moyen d'un modèle "hors-ligne" permettant de résoudre une chimie complexe le long de lignes de courant extraites d'un écoulement moyenné dans le temps résultant du calcul précédent (non réactif). Enfin, une SGE multi-espèces est réalisée, incluant un schéma chimique auparavant réduit afin de limiter le coût de calcul. Cette simulation représente une des toutes premières SGE d'un jet supersonique réactif, incluant la tuyère, effectuée sur un domaine de calcul aussi long. En capturant avec précision le mélange du panache avec l'air ambiant ainsi que les interactions entre turbulence et combustion, la technique des simulations aux grandes échelles offre une évaluation des concentrations des espèces chimiques dans le jet d'une precision inédite. Ces résultats peuvent être utilisés pour initialiser des calculs atmosphériques sur de plus larges domaines, afin de modéliser les réactions entre chlore actif et ozone et de quantifier l'appauvrissement en ozone dans le panacheRockets have an impact on the chemical composition of the atmosphere, and particularly on stratospheric ozone. Among all types of propulsion, Solid-Rocket Motors (SRMs) have given rise to concerns since their emissions are responsible for a severe decrease in ozone concentration in the rocket plume during the first hours after a launch. The main source of ozone depletion is due to the conversion of hydrogen chloride, a chemical compound emitted in large quantities by ammonium perchlorate based propellants, into active chlorine compounds, which then react with ozone in a destructive catalytic cycle, similar to those responsible for the Antartic "Ozone hole". This conversion occurs in the hot, supersonic exhaust plume, as part of a strong second combustion between chemical species of the plume and air. The objective of this study is to evaluate the active chlorine concentration in the far-field plume of a solid-rocket motor using large-eddy simulations (LES). The gas is injected through the entire nozzle of the SRM and a local time-stepping method based on coupling multi-instances of the fluid solver is used to extend the computational domain up to 400 nozzle exit diameters downstream of the nozzle exit. The methodology is validated for a non-reactive case by analyzing the flow characteristics of the resulting supersonic co-flowing under-expanded jet. Then the chemistry of chlorine is studied off-line using a complex chemistry solver applied on trajectories extracted from the LES time-averaged flow-field. Finally, the online chemistry is analyzed by means of the multi-species version of the LES solver using a reduced chemical scheme. To the best of our knowledge, this represents one of the first LES of a reactive supersonic jet, including nozzle geometry, performed over such a long computational domain. By capturing the effect of mixing of the exhaust plume with ambient air and the interactions between turbulence and combustion, LES offers an evaluation of chemical species distribution in the SRM plume with an unprecedented accuracy. These results can be used to initialize atmospheric simulations on larger domains, in order to model the chemical reactions between active chlorine and ozone and to quantify the ozone loss in SRM plumes
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Heiner, Mark C. "Development and Testing of a Hydrogen Peroxide Injected Thrust Augmenting Nozzle for a Hybrid Rocket." DigitalCommons@USU, 2019. https://digitalcommons.usu.edu/etd/7630.
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During a rocket launch, the point at which the most thrust is needed is at lift-off where the rocket is the heaviest since it is full of propellant. Unfortunately, this is also the point at which rocket engines perform the most poorly due to the relatively high atmospheric pressure at sea level. The Thrust Augmenting Nozzle (TAN) investigated in this paper provides a solution to this dilemma. By injecting extra propellant into the nozzle but downstream of the throat, the internal nozzle pressure is raised and the thrust is increased, and the nozzle efficiency, or specific impulse is potentially improved as well. Using this concept, the payload capacity of a launch vehicle can be increased and provides an excellent option for single stage to orbit vehicles.
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Ingenito, Antonella. "Modellistica della combustione in regime supersonico." Doctoral thesis, La Sapienza, 2006. http://hdl.handle.net/11573/917117.
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Stoukov, Alexei. "Etude numérique de la couche de mélange réactive supersonique." Rouen, 1996. http://www.theses.fr/1996ROUES013.
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L'objet de ce travail concerne l'analyse et la simulation numérique de la couche de mélange supersonique réactive instationnaire. Le premier chapitre présente le contexte général de l'étude et les problèmes spécifiques abordés. Le modèle physico-chimique comprenant les équations de Navier-Stokes et la cinétique chimique complexe est présenté dans la suite. Le troisième chapitre consiste en une étude comparative de différents schémas numériques de résolution des problèmes de type hyperbolique et présente une validation du code numérique développé autour du schéma TVD Upwind de Harten-Yee. Le traitement numérique des conditions aux limites et plus particulièrement des conditions de non-réflexion est présenté dans le chapitre suivant. Le reste de ce mémoire est consacré à l'interprétation de résultats de calculs de couches de mélange air-hydrogène (dans un premier temps inertes, puis réactives). Dans ce chapitre, une attention particulière a été portée sur l'étude phénoménologique du processus de mélange conditionné par les structures à grandes échelles. Par la suite les problèmes liés à l'interaction de ces structures avec onde de choc oblique sont abordés ; l'influence de ces structures sur le processus d'auto-allumage et sur la flamme qui en résulte est finalement analysée.
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Jeyashekar, Nigil Satish. "Temperature and number density measurements using Raman scattering in turbulent-supersonic-combusting flows /." Full text available from ProQuest UM Digital Dissertations, 2006. http://0-proquest.umi.com.umiss.lib.olemiss.edu/pqdweb?index=0&did=1379528381&SrchMode=1&sid=4&Fmt=2&VInst=PROD&VType=PQD&RQT=309&VName=PQD&TS=1217357303&clientId=22256.
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Santos, Wagner Matos. "Diminuição de rejeição térmica em motores a combustão /." Guaratinguetá, 2019. http://hdl.handle.net/11449/192781.
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Orientador: Marcelino Pereira do NascimentoResumo: Estudos recentes mostram tendência de desenvolvimento de sistemas capazes de recuperar energia dos gases de escape. Apesar de ser de fato uma fonte de energia, os sistemas para recuperação instalados na linha de exaustão mostram grande complexidade e necessidade de muito espaço de instalação. O presente estudo demonstra a viabilidade de recuperação de energia do sistema de arrefecimento com espaço de instalação reduzido. O estudo propõe a geração de vapor através de perfis instalados nas paredes da câmara de combustão e sua aplicação em um ciclo Rankine para gerar trabalho através de expansor. A simulação realizada com o software GT-Power demonstra a possibilidade de extrair calor das câmaras com a mesma eficiência de um sistema convencional de arrefecimento por água e aditivo. Paralelamente, o estudo propõe um sistema de recirculação de gases de exaustão sem o uso de um trocador de calor, patenteado pela empresa patrocinadora, complementando a estratégia para evitar o uso desnecessário de energia em um motor a combustão. Os gases provenientes da linha de baixa pressão da exaustão são admitidos pela linha de ar limpo, em alta pressão, através do uso de uma válvula Laval. Os resultados são provenientes de uma rotina de cálculo em Excel, base VBA. O sistema permite que os gases fluam por fora da turbina e elimina a necessidade de um trocador de calor, resolvendo o maior problema dos sistema convencionais de recirculação de gases em baixa pressão. A combinação das duas soluções,... (Resumo completo, clicar acesso eletrônico abaixo)
Doutor
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Lee, Hsiyu-Fu, and 李旭富. "Investigation on Supersonic Combustion Ramjet Engine." Thesis, 1990. http://ndltd.ncl.edu.tw/handle/94040094870209702042.
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碩士淡江大學
機械工程研究所
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Supersonic Combustion Ramjet (SCRAMJET) engine shall be the primary propulsion system for aerospace airplanes in the future and be the latest airbreathing engine. This thesis is to make cycle analysis of each of its components and to make some modification of individual component analysis from the thesis proposed by 0' Yang[17]. Finally, it makes analytical processes of components to be more completely corresponding to one-dimensional integral approach. This thesis takes the theory of a unified cycle analysis and discusses the performance of the SCRAMJET engine. The combining relations of SCRAMJET engine and integral-airframe of aerospace plane is also be concerned. The profile of specific impluse with respect to flight mach number is satisfactory from the analysis of this study. Also, it proves the excellence of the SCRAMJET engine.
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Joarder, Ratan. "Demonstration Of Supersonic Combustion In A Combustion Driven Shock-Tunnel." Thesis, 2009. http://hdl.handle.net/2005/1005.
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For flights beyond Mach 6 ramjets are inefficient engines due to huge total pressure loss in the normal shock systems, combustion conditions that lose a large fraction of the available chemical energy due to dissociation and high structural loads. However if the flow remains supersonic inside the combustion chamber, the above problems could be alleviated and here the concept of SCRAMJET(supersonic combustion ramjet) comes into existence. The scramjets could reduce launching cost of satellites by carrying only fuel and ingesting oxygen from atmospheric air. Further applications could involve defense and transcontinental hypersonic transport.
In the current study an effort is made to achieve supersonic combustion in a ground based short duration test facility(combustion driven shock-tunnel), which in addition to flight Mach number can simulate flight Reynolds number as well. In this study a simple method of injection i.e. wall injection of the fuel into the combustion chamber is used. The work starts with threedimensional numerical simulation of a non-reacting gas(air) injection into a hypersonic cross-flow of air to determine the conditions in which air penetrates reasonably well into the cross-flow. Care is taken so that the process does not induce huge pressure loss due to the bow shock which appears in front of the jet column. The code is developed in-house and parallelized using OpenMp model. This is followed by experiments on air injection into a hypersonic cross-flow of air in a conventional shock-tunnel HST2 existing in IISc. The most tricky part is synchronization of injection with start of test-flow in such a short duration(test time 1 millisecond) facility.
Next part focuses on numerical simulations to determine the free-stream conditions, mainly the temperature and pressure of air, so that combustion takes place when hydrogen is injected into a supersonic cross-flow of air. The simulations are two-dimensional and includes species conservation equations and source terms due to chemical reactions in addition to the Navier-Stokes equations. This code is also built in-house and parallelized because of more number of operations with the inclusion of species conservation equations and chemical non-equilibrium. However, the predicted conditions were not achievable by HST2 due to low stagnation conditions of HST2.
Therefore, a new shock-tunnel which could produce the required conditions is built. The new tunnel is a combustion driven shock-tunnel in which the driver gas is at higher temperature than conventional shock-tunnel. The driver gas is basically a mixture of hydrogen, oxygen and helium at a mole ratio of 2:1:10 initially. The mixture is ignited by spark plugs and the hydrogen and oxygen reacts releasing heat. The heat released raises the temperature of the mixture which is now predominantly helium and small fractions of water vapour and some radicals. The composition of the driver gas and initial pressure are determined through numerical simulations.
Experiments follow in the new tunnel on hydrogen injection into a region of supersonic cross-flow between two parallel plates with a wedge attached to the bottom plate. The wedge reduces the hypersonic free-stream to Mach 2. A high-speed camera monitors the flow domain around injection point and sharp rise in luminosity is observed. To ascertain whether the luminosity is due to combustion or not, two more driven gases namely nitrogen and oxygen-rich air are used and the luminosity is compared. In the first case, the free-stream contains no oxygen and luminosity is not observed whereas in the second case higher luminosity than air driver case is visible. Additionally heat-transfer rates are measured at the downstream end of the model and at a height midway between the plates. Similar trend is observed in the relative heat-transfer rates. Wall static pressure at a location downstream of injection port is also measured and compared with numerical simulations. Results of numerical simulations which are carried out at the same conditions as of experiments confirm combustion at supersonic speed.
Experiments and numerical simulations show presence of supersonic combustion in the setup. However, further study is necessary to optimize the parameters so that thrust force could be generated efficiently.