Dissertations / Theses: 'Scramjets'

1

Hirschen, Christian. "Experimentelle Untersuchungen zur Düsen- und Heckströmung eines Scramjets." Köln DLR, Bibliotheks- und Informationswesen, 2009. http://d-nb.info/1000055507/34.

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2

Tsimis, Charalampos. "Fuel jet injection and supersonic mixing for scramjets." Thesis, Imperial College London, 2007. http://hdl.handle.net/10044/1/8722.

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3

Tran, Kathleen. "One Dimensional Analysis Program for Scramjet and Ramjet Flowpaths." Thesis, Virginia Tech, 2010. http://hdl.handle.net/10919/30857.

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One-Dimensional modeling of dual mode scramjet and ramjet flowpaths is a useful tool for scramjet conceptual design and wind tunnel testing. In this thesis, modeling tools that enable detailed analysis of the flow physics within the combustor are developed as part of a new one-dimensional MATLAB-based model named VTMODEL. VTMODEL divides a ramjet or scramjet flow path into four major components: inlet, isolator, combustor, and nozzle. The inlet module provides two options for supersonic inlet one-dimensional calculations; a correlation from MIL Spec 5007D, and a kinetic energy efficiency correlation. The kinetic energy efficiency correlation also enables the user to account for inlet heat transfer using a total temperature term in the equation for pressure recovery. The isolator model also provides two options for calculating the pressure rise and the isolator shock train. The first model is a combined Fanno flow and oblique shock system. The second model is a rectangular shock train correlation. The combustor module has two options for the user in regards to combustion calculations. The first option is an equilibrium calculation with a â growing combustion sphereâ combustion efficiency model, which can be used with any fuel. The second option is a non-equilibrium reduced-order hydrogen calculation which involves a mixing correlation based on Mach number and distance from the fuel injectors. This model is only usable for analysis of combustion with hydrogen fuel. Using the combustion reaction models, the combustor flow model calculates changes in Mach number and flow properties due to the combustion process and area change, using an influence coefficient method. This method iii also can take into account heat transfer, change in specific heat ratio, change in enthalpy, and other thermodynamic properties. The thesis provides a description of the flow models that were assembled to create VTMODEL. In calculated examples, flow predictions from VTMODEL were compared with experimental data obtained in the University of Virginia supersonic combustion wind tunnel, and with reported results from the scramjet models SSCREAM and RJPA. Results compared well with the experiment and models, and showed the capabilities provided by VTMODEL.
Master of Science

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4

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 cavité
The 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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5

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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6

Maddalena, Luca. "Investigations of Injectors for Scramjet Engines." Diss., Virginia Tech, 2007. http://hdl.handle.net/10919/28683.

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An experimental study of an aerodynamic ramp (aeroramp) injector was conducted at Virginia Tech. The aeroramp consisted of an array of two rows with two columns of flush-wall holes that induce vorticity and enhance mixing. For comparison, a single-hole circular injector with the same area angled downstream at 30 degrees was also examined. Test conditions involved sonic injection of helium heated to 313 K, to safely simulate hydrogen into a Mach 4 air cross-stream with average Reynolds number 5.77 e+7 per meter at a jet to freestream momentum flux ratio of 2.1. Sampling probe measurements were utilized to determine the local helium concentration. Pitot and cone-static pressure probes and a diffuser thermocouple probe were employed to document the flow. The main results of this work was that the mixing efficiency value of this aeroramp design which was optimized at Mach 2.4 for hydrocarbon fuel was only slightly higher than that of the single-hole injector at these flow conditions and the mass-averaged total pressure loss parameter showed that the aero-ramp and single-hole injectors had the same overall losses. The natural extension of the investigation was then to look in detail at two major physical phenomena that occurs in a complex injector design such the Aeroramp: the jet-shock interaction and the interaction of the vortical structures produced by the jets injection into a supersonic cross flow. Experimental studies were performed to investigate the effects of impinging shocks on injection of heated helium into a Mach 4 crossflow. It was found that the addition of a shock behind gaseous injection into a Mach 4 crossflow enhances mixing only if the shock is closer to the injection point where the counter-rotating vortex pair (always associated with transverse injection in a crossflow) is not yet formed, and the deposition of baroclinic generated of vorticity is the highest. The final investigation concerned with the interaction of the usual vortex structure produced by jet injection into a supersonic crossflow and an additional axial vortex typical of those that might be produced by the inlet of a scramjet or the forebody of a vehicle to be controlled by jet interaction phenomena. The additional axial vortices were generated by a strut-mounted, diamond cross-section wing mounted upstream of the injection location. The wing was designed to produce a tip vortex of a strength comparable to that of one of the typical counter-rotating vortex pair (CVP) found in the plume of a jet in a crossflow. The profound interaction of supersonic vortices supported by a quantitative description and characterization of the flowfield has been demonstrated.
Ph. D.

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7

Anderson, Cody Dean. "Development and Testing of an Integrated Liquid-Fuel-Injector/Plasma-Igniter for Scramjets." Thesis, Virginia Tech, 2004. http://hdl.handle.net/10919/31416.

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A newly designed liquid fuel (kerosene) aeroramp injector/plasma igniter was tested in cold flow using the Virginia Tech supersonic wind tunnel at Mach 2.4. The liquid fuel (kerosene) injector is flush wall mounted and consists of a 2 hole aeroramp array of impinging jets that are oriented in a manner to improve mixing and atomization of the liquid jets. The two jets are angled downstream at 40 degrees and have a toe-in angle of 60 degrees. The plasma torch used nitrogen and air as feedstocks and was placed downstream of the injector as an ignition aid. First, schlieren and shadowgraph photographs were taken of the injector flow to study the behavior of the jets, shape of the plume, and penetration of the liquid jet. The liquid fuel aeroramp was found to have better penetration than a single, round jet at 40 degrees. However, the liquid fuel aeroramp does not penetrate as well as an upstream/downstream impinging jet in a plane aligned with the flow. Next, the Sauter mean droplet diameter distribution was measured downstream of the injector. The droplet diameter was found to vary from 21 to 37 microns and the atomization of the injector does not appear to improve beyond 90 effective jet diameters from the liquid fuel aeroramp. These results were then used to decide on an initial location for the plasma torch. The combined liquid injector/plasma torch system was tested in an unheated (300 K) Mach 2.4 flow with a total pressure of 345 kPa. The liquid fuel (kerosene) volumetric flow rate was varied from 0.66 lpm to 1.22 lpm for the combined liquid injector/plasma torch system. During this testing the plasma torch was operated from 1000 to 5000 watts with 25 slpm of nitrogen and air as feedstocks. The interaction between the spray plume and the plasma torch was observed with direct photographs, videos, and photographs through an OH filter. It is difficult to say that any combustion is present from these photographs. Of course, it would be surprising if much combustion did occur under these cold-flow, low-pressure conditions. Differences between the interaction of the spray plume and the plasma torch with nitrogen and air as feedstocks were documented. According to the OH wavelength filtered photographs the liquid fuel flow rate does appear to have an effect on the height and width of the bright plume. As the liquid fuel flow rate increases the bright plume increases in height by 30% and increases in width slightly (2%). While, a decrease in liquid fuel flow rate resulted in an increase in height by 9% and an increase in width by 10%. Thus, as the liquid fuel flow rate varies the width and height of the bright plume appear to always increase. This can be explained by noticing that the shape of the bright plume changes as the liquid fuel flow rate varies and perhaps anode erosion during testing also plays a part in this variation of the bright plume. From the OH wavelength filtered photographs it was also shown that the bright plume appears to decrease in width by 9% and increase in height by 22% when the plasma torch is set at a lower power setting. When air is used as the torch feedstock, instead of nitrogen, the penetration of the bright plume can increase by as much as 19% in width and 17% in height. It was also found that the height and width of the bright plume decreased slightly (2%) as the fuel flow rate increased when using air as the torch feedstock. Testing in a hot-flow facility is planned.
Master of Science

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8

Grossman, Peter Michael. "Experimental Investigation of a Flush-Walled, Diamond-Shaped Fuel Injector for High Mach Number Scramjets." Thesis, Virginia Tech, 2007. http://hdl.handle.net/10919/30974.

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An experimental investigation of a flush-wall, diamond-shaped injector was conducted in the Virginia Tech supersonic wind tunnel. The diamond injector was elongated in the streamwise direction and is aimed downstream angled up at 60° from the wall. Test conditions involved sonic injection of helium heated to approximately 313 K into a nominal Mach 4.0 crossstream airflow. These conditions are typical of a scramjet engine for a Mach 10 flight, and heated helium was used to safely simulate hydrogen fuel. The injector was tested at two different injectant conditions. First, it was investigated at a baseline mass flow rate of 3.4 g/s corresponding to an effective radius of 3.54 mm and a jet-to-freestream momentum flux ratio of 1.04. Second, a lower mass flow rate of 1.5 g/s corresponding to an effective ratio of 2.35 mm and a jet-to-freestream momentum flux ratio of 0.49 was studied. The diamond injector was tested both aligned with the freestream and at a 15° yaw angle for the baseline mass flow rate and aligned with the freestream at the lower mass flow rate. For comparison, round injectors angled up at 30° from the wall were also examined at both flow rates. A smaller round injector was used at the lower mass flow rate such that the jet-to-freestream momentum flux ratio was 1.75 for both cases. A concentration sampling probe and gas analyzer were used to determine the local helium concentration, while Pitot, cone-static and total temperature probes were used to determine the flow properties.

The results of the investigation can be summarized as follows. For the baseline case, the aligned diamond injector penetrated 44% higher into the crossflow than did the round injector. The addition of yaw angle increased the crossflow penetration to 53% higher than the round injector. The aligned diamond injector produced a 34% wider jet than the round injector, while the addition of yaw angle somewhat reduced this widening effect to 26% wider than the round injector. The aligned and yawed diamond injectors exhibited 10% and 15% lower mixing efficiency than the round injector, respectively. The total pressure loss parameter of the aligned diamond was 22% lower than the round injector, while the addition of yaw angle improved the total pressure loss parameter to 34% lower than the round injector. For the lower mass flow (and momentum flux ratio) case, the diamond injector demonstrated 52% higher penetration and a 39% wider plume than the round injector. The mixing efficiency was nearly identical between the two injectors with just a 4% lower mixing efficiency for the diamond injector. The total pressure loss parameter of the diamond injector was 32% lower than round injector. These results confirm the conclusions of earlier, lower free stream Mach number and higher molecular weight injectant, studies that a slender diamond injector provides significant benefits for crossflow penetration and lower total pressure losses.
Master of Science

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9

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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10

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 10Âº 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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11

Eberhardt, Sebastian [Verfasser], Nikolaus A. [Akademischer Betreuer] [Gutachter] Adams, and Stefan [Gutachter] Hickel. "Large-eddy simulations of supersonic mixing with application to scramjets / Sebastian Eberhardt ; Gutachter: Stefan Hickel, Nikolaus A. Adams ; Betreuer: Nikolaus A. Adams." München : Universitätsbibliothek der TU München, 2017. http://d-nb.info/1130786544/34.

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12

Hirschen, Christian [Verfasser]. "Experimentelle Untersuchungen zur Düsen- und Heckströmung eines Scramjets / Christian Hirschenhristian Hirschen. DLR, Deutsches Zentrum für Luft- und Raumfahrt e.V. in der Helmholtz-Gemeinschaft. [Hrsg. Deutsches Zentrum für Luft- und Raumfahrt e.V., Bibliotheks- und Informationswesen]." Köln : DLR, Bibliotheks- und Informationswesen, 2009. http://d-nb.info/1000055507/34.

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13

Matthews, Alexander J. "Scramjet intakes." Thesis, University of Oxford, 2003. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.400217.

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14

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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15

Schütte, Gerrit. "Probabilistische Untersuchungen zu Scramjet-Antriebssystemen." Aachen Shaker, 2009. http://d-nb.info/1000474291/04.

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16

McGuire, Jeffrey Robert Aerospace Civil &amp 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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17

Biasca, Rodger Joseph. "Chemical kinetics of SCRAMJET propulsion." Thesis, Massachusetts Institute of Technology, 1988. http://hdl.handle.net/1721.1/35949.

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18

Ingenito, Antonella. "Modellistica della combustione in regime supersonico." Doctoral thesis, La Sapienza, 2006. http://hdl.handle.net/11573/917117.

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19

Gardner, Anthony D. "HyShot scramjet testing in the HEG /." Köln : DLR, Bibliotheks- und Informationswesen, 2007. http://bvbr.bib-bvb.de:8991/F?func=service&doc_library=BVB01&doc_number=016271138&line_number=0001&func_code=DB_RECORDS&service_type=MEDIA.

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20

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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21

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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22

Brandstetter, Armin Johann Leopold. "Betriebsverhalten einer Dualmodus-SCRamjet-Modellbrennkammer mit Wasserstoffverbrennung." [S.l. : s.n.], 2004. http://deposit.ddb.de/cgi-bin/dokserv?idn=974166103.

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23

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, 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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24

Munro, Stuart. "Scramjet Intakes: Designing for Performance and Operability." Thesis, University of Oxford, 2007. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.491519.

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The classical approach to scramjet intake design, analysis, experimental testing, and performance assessment treats the intake as separate from the rest of the vehicle. This thesis argues that a more integrated approach is necessary. The design space of Mach 7 two-dimensional, axi-symmetric spike and Busemann-type intakes with focused forebody waves and throats expanded to the Kantrowitz limit is parameterised geometrically and stream thrust average performance parameters calculated across the range of possible intake designs. An engine cycle is applied to arbitrate between conflicting capability and efficiency design points. The two-dimensional analysis is extended to include angle of attack, off-design Mach number and defocused conditions. Intakes with sharp leading edges and moderate turning angles are found to perform optimally, and the relatively large combustor hydraulic diameter and higher capability of inward turning designs are found to improve engine performance significantly. An upper bound is established for the performance of stream-traced Busemann intakes and it is shown how the performance of truncated Busemann derivatives may be improved by extending the leading edge at constant angle. The thesis also describes complementary experiments investigating intake operability carried out in the University of Oxford's gun tunnel. This facility was converted to operate as a Ludwieg tube with Light piston Isentropic Compression Heating (LICH) to improve flow steadiness when matching Reynolds numbers and temperature ratios to 2:5 scale Mach 7, 30 km altitude flight. Tunnel stagnation temperatures are measured and correlations to the fill pressures and stagnation pressure history are determined. The design and build of a two-dimensional wedge intake with isolator to determine intake starting limits by varying cowl position and shoulder angle is described. The isolator is shown to degrade starting characteristics by increasing the likelihood of internal choking. The implementation of a self-starting test, employing gas injection via fast acting valves, is implemented to demonstrate the self-starting characteristics of an axi-symmetric spike intake with internal contraction exceeding the Kantrowitz limit. Fuel injection and combustor back pressure are also simulated at angle of attack. Boundary layer trips are shown to significantly improve the operability of both models tested.

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25

Schütte, Gerrit [Verfasser]. "Probabilistische Untersuchungen zu Scramjet-Antriebssystemen / Gerrit Schütte." Aachen : Shaker, 2010. http://d-nb.info/1124364838/34.

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26

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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27

Halter, Megaera C. "Energy-turns analysis for a scramjet powered missle." Thesis, Virginia Tech, 1988. http://hdl.handle.net/10919/43753.

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A reduced order model describing the energy and heading angle dynamics of a scramjet missile is developed using a singular perturbation technique. The cruise analysis is briefly reviewed to determine the conditions at which the missile will cruise most efficiently. The turn and climb performance of the missile over the conditions of interest is then examined and a family of extremal trajectories is constructed which asymptotically approach the cruise at an intermediate altitude.
Master of Science

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28

Barone, Dominic L. "Investigation of TDLAS Measurements in a Scramjet Engine." University of Cincinnati / OhioLINK, 2010. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1277130335.

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29

Omari, Dennis. "Control of Transient Unstart in Isolator of Scramjet." The Ohio State University, 2018. http://rave.ohiolink.edu/etdc/view?acc_num=osu1523367091946973.

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30

Frost, Myles Alexander. "Hyshot scramjet experiments in the T4 shock tunnel /." [St. Lucia, Qld.], 2001. http://www.library.uq.edu.au/pdfserve.php?image=thesisabs/absthe16919.pdf.

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31

Fuhrmann, Thomas. "Auslegung und Betriebsverhalten von SCRamjet-Antriebssystemen für Raumtransporter-Hyperschallflugzeuge." München Verl. Dr. Hut, 2009. http://mediatum2.ub.tum.de/node?id=735781.

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32

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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33

Brewer, Keith Merritt. "Exergy Methods for the Mission-Level Analysis and Optimization of Generic Hypersonic Vehicles." Thesis, Virginia Tech, 2006. http://hdl.handle.net/10919/32007.

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Though the field of hypersonic vehicle design is thriving again, few studies to date demonstrate the technology through a mission in which multiple flight conditions and constraints are encountered. This is likely due to the highly integrated and sensitive nature of hypersonic vehicle components. Consequently, a formal Mach 6 through Mach 10 flight envelope is explored which includes cruise, acceleration/climb, deceleration/descend and turn mission segments. An exergy approach to the vehicle synthesis/design, in which trade-offs between dissimilar technologies are observed, is proposed and measured against traditional methods of assessing highly integrated systems. A quasi one-dimensional hypersonic vehicle system simulation program was constructed. Composed of two sub-systems, propulsion and airframe, mechanisms for loss are computed from such irreversible processes as shocks, friction, heat transfer, mixing, and incomplete combustion. The propulsion sub-system consists of inlet, combustor, and nozzle, while the airframe provides trim and force accounting measures. An energy addition mechanism, based on the potential of MHD technology, is utilized to maintain a shock-on-lip inlet operating condition. Thirteen decision variables (seven design and six operational) were chosen to govern the vehicle geometry and performance. A genetic algorithm was used to evaluate the optimal vehicle synthesis/design for three separate objective functions, i.e the optimizations involved the maximization of thrust efficiency, the minimization of fuel mass consumption, and the minimization of exergy destruction plus fuel exergy loss. The principal results found the minimum fuel consumption and minimum exergy destruction measures equivalent, both meeting the constraints of the mission while using 11% less fuel than the thrust efficiency measure. Optimizing the vehicle for the single most constrained mission segment yielded a vehicle capable of flying the entire mission but with fuel consumption and exergy destruction plus fuel loss values greater than the above mentioned integrated vehicle solutions. In essence, the mission-level analysis provided much insight into the dynamics of mission-level hypersonic flight and demonstrated the usefulness of an exergy destruction minimization measure for highly integrated synthesis/design.
Master of Science

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34

Markell, Kyle Charles. "Exergy Methods for the Generic Analysis and Optimization of Hypersonic Vehicle Concepts." Thesis, Virginia Tech, 2005. http://hdl.handle.net/10919/31256.

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This thesis work presents detailed results of the application of exergy-based methods to highly dynamic, integrated aerospace systems such as hypersonic vehicle concepts. In particular, an exergy-based methodology is compared to a more traditional based measure by applying both to the synthesis/design and operational optimization of a hypersonic vehicle configuration comprised of an airframe sub-system and a propulsion sub-system consisting of inlet, combustor, and nozzle components. A number of key design and operational decision variables are identified as those which govern the hypersonic vehicle flow physics and thermodynamics and detailed one-dimensional models of each component and sub-system are developed. Rates of exergy loss as well as exergy destruction resulting from irreversible loss mechanisms are determined in each of the hypersonic vehicle sub-systems and their respective components. Multiple optimizations are performed for both the propulsion sub-system only and for the entire hypersonic vehicle system for single mission segments and for a partial, three-segment mission. Three different objective functions are utilized in these optimizations with the specific goal of comparing exergy methods to a standard vehicle performance measure, namely, the vehicle overall efficiency. Results of these optimizations show that the exergy method presented here performs well when compared to the standard performance measure and, in a number of cases, leads to more optimal syntheses/designs in terms of the fuel mass flow rate required for a given task (e.g., for a fixed-thrust requirement or a given mission). In addition to the various vehicle design optimizations, operational optimizations are conducted to examine the advantage if any of energy exchange to maintain shock-on-lip for both design and off-design conditions. Parametric studies of the hypersonic vehicle sub-systems and components are also conducted and provide further insights into the impacts that the design and operational decision variables and flow properties have on the rates of exergy destruction.
Master of Science

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35

Fuhrmann, Sirka [Verfasser]. "Gestufte Injektion und Verbrennung in einer Scramjet-Brennkammer / Sirka Fuhrmann." München : Verlag Dr. Hut, 2013. http://d-nb.info/1033041815/34.

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36

Frauholz, Sarah [Verfasser]. "Advanced Numerical Investigations of Hypersonic Scramjet Intake Flows / Sarah Frauholz." München : Verlag Dr. Hut, 2015. http://d-nb.info/107406318X/34.

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37

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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38

Zinnecker, Alicia M. "Modeling for Control Design of an Axisymmetric Scramjet Engine Isolator." The Ohio State University, 2012. http://rave.ohiolink.edu/etdc/view?acc_num=osu1354215841.

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39

Larios-Barbosa, Jaime Omar. "SHOCK CORRELATION INVESTIGATION IN A GASEOUS FUELED AXISYMMETRIC SCRAMJET FLOWPATH." Wright State University / OhioLINK, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=wright1377130377.

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40

Stewart, Benjamin S. "Predicted scramjet testing capabilities of the proposed RHYFL-X expansion tube /." [St. Lucia, Qld], 2004. http://www.library.uq.edu.au/pdfserve.php?image=thesisabs/absthe18241.pdf.

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41

Fischer, Christian Max [Verfasser]. "Investigation of the isolator flow of scramjet engines / Christian Max Fischer." Aachen : Hochschulbibliothek der Rheinisch-Westfälischen Technischen Hochschule Aachen, 2014. http://d-nb.info/1059796627/34.

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42

Krause, Martin [Verfasser]. "Numerical Analysis of Transition Effects for SCRamjet Intake Flows / Martin Krause." Aachen : Shaker, 2010. http://d-nb.info/1101185546/34.

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43

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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44

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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45

Nguyen, Tran Quang Tue [Verfasser]. "Numerical investigations of relaminarization in scramjet flows / Tran Quang Tue Nguyen." Aachen : Hochschulbibliothek der Rheinisch-Westfälischen Technischen Hochschule Aachen, 2012. http://d-nb.info/1023093197/34.

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46

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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47

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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48

Burger, Scott Kuhlman. "Investigation of Injectant Molecular Weight and Shock Impingement Effects on Transverse Injection Mixing in Supersonic Flow." Thesis, Virginia Tech, 2010. http://hdl.handle.net/10919/31981.

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This study examines the effect of varying injectant molecular weight on the penetration of transverse injection jets into a supersonic crossflow. The injectants considered here are methane (W=16.04), air (W=28.97) and carbon dioxide, (W=44.01). These results augment the previous results obtained at Virginia Tech for helium (W=4.00) injection under the same test conditions to provide a very wide range of molecular weights. Second, since shocks are ubiquitous in scramjet combustors, their influence on penetration and mixing was also studied by arranging for an oblique shock to impinge near the injection station. The cases of a shock impinging upstream and downstream of the injector were both examined. One can anticipate an important influence of molecular weight here also because of the importance of density gradients on the generation of vorticity by baroclinic torque. Increasing molecular weight was found to increase penetration in general, as well as increase the lateral spreading of the plume. The majority of the data shows a weak dependency of the jet size on molecular weight, but there are indications that under certain circumstances large changes in the flow structure may occur due to molecular weight effects. The addition of an impinging shock is found to increase mixing and decrease penetration and plume size, especially with the shock impinging downstream of the injector.
Master of Science

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49

Cox-Stouffer, Susan K. Jr. "Numerical Simulation of Injection and Mixing in Supersonic Flow." Diss., Virginia Tech, 1997. http://hdl.handle.net/10919/29628.

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A numerical investigation of the performance of two candidate designs for injection into supersonic flow, including a comparison of two renormalized group theory (RNG) based k-epsilon turbulence models with a more conventional k-epsilon model. The chosen designs were an unswept ramp injector with four injection ports and a novel nine-hole injector array. The objectives of the investigation were to provide reliable computational solutions to the flowfields in question using both RNG and standard k-epsilon turbulence models and to compare the solutions to experiment, thereby to judge the relative performance of the turbulence models. A second objective of the investigation was to use the computed data to provide design insights for the nine-hole injector array. This investigation made use of GASP(tm) version 2.2, a commercial computational fluid dynamics code that was augmented by the addition of one RNG-based k-epsilon turbulence model derived by Zhou, et. al. and one variant of Zhou's model, which was derived by the author. Mesh sequencing studies were performed to measure solution quality, with the fine mesh for the injector array containing roughly one million grid nodes and the fine mesh for the ramp injector containing more than six million grid nodes. Results of these studies indicated that the injector-array solution was significantly under-resolved in the farfield, though the quality was better in the vicinity of the injector itself. The ramp-injector solution, while not perfectly grid-resolved, showed much better grid convergence in both the nearfield and farfield. Accordingly, comparison with experiment was better for the ramp injector than for the injector array. For both injectors, the differences between solutions generated with RNG-based k-epsilon and standard k-epsilon turbulence models were negligibly small." Despite inadequate grid resolution in the farfield, the computational investigation of the nine-hole injector array did yield several important design insights. Particularly, the significance to mixing and losses of the placement of the outer injectors of the second and third rows was determined.
Ph. D.

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50

Robinson, Matthew J. "Simultaneous lift, moment and thrust measurement on a scramjet in hypervelocity flow /." [St. Lucia, Qld.], 2003. http://www.library.uq.edu.au/pdfserve.php?image=thesisabs/absthe17611.pdf.

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Dissertations / Theses on the topic 'Scramjets'

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