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1
Goodman, J. S. "Thermal analysis of ramjet engines." Thesis, University of Oxford, 2007. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.445768.
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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.
3
Reuter, Dierk Martin. "Investigation of combustion instability in ramjet combustors." Diss., Georgia Institute of Technology, 1988. http://hdl.handle.net/1853/12271.
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Davis, James Arthur. "Acoustic-vortical-combustion interaction in a solid fuel ramjet simulator." Diss., Georgia Institute of Technology, 1989. http://hdl.handle.net/1853/12947.
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Hall, Philip D. "Design of a coaxial split flow pulse detonation engine." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2006. http://library.nps.navy.mil/uhtbin/hyperion/06Jun%5FHall.pdf.
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Thesis (M.S. in Mechanical Engineering)--Naval Postgraduate School, June 2006.Thesis Advisor(s): Jose O. Sinibaldi, Christopher M. Brophy. "June 2006." Includes bibliographical references (p. 41-42). Also available in print.
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De, Groot Wim A. (Wim Adrianus). "Laser Doppler diagnostics of the flow behind a backward facing step." Diss., Georgia Institute of Technology, 1985. http://hdl.handle.net/1853/15801.
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Najafiyazdi, Alireza. "Theoretical and numerical analysis of supersonic inlet starting by mass spillage." Thesis, McGill University, 2007. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=111524.
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Supersonic inlet starting by mass spillage is studied theoretically and numerically in the present thesis. A quasi-one-dimensional, quasi-steady theory is developed for the analysis of flow inside a perforated inlet. The theory results in closed-form relations applicable to flow starting by the mass spillage technique in supersonic and hypersonic inlets.The theory involves three parameters to incorporate the multi-dimensional nature of mass spillage through a wall perforation. Mass spillage through an individual slot is studied to determine these parameters; analytical expressions for these parameters are derived for both subsonic and supersonic flow conditions. In the case of mass spillage from supersonic flows, the relations are exact. However, due to the complexity of flow field, the theory is an approximation for subsonic flows. Therefore, a correction factor is introduced which is determined from an empirical relation obtained from numerical simulations.
A methodology is also proposed to determine perforation size and distribution to achieve flow starting for a given inlet at a desired free-stream Mach number. The problem of shock stability inside a perforated inlet designed with the proposed method is also discussed.
The method is demonstrated for some test cases. Time-realistic CFD simulations and experimental results in the literature confirm the accuracy of the theory and the reliability of the proposed design methodology.
8
Sarisin, Mustafa Nevzat. "Design Of A Connected Pipe Test Facility For Ramjet Applications." Master's thesis, METU, 2005. http://etd.lib.metu.edu.tr/upload/12606078/index.pdf.
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ABSTRACT
DESIGN OF A CONNECTED PIPE TEST FACILITY
FOR RAMJET APPLICATIONS
SARISIN, Mustafa Nevzat
M.S., Department of Mechanical Engineering
Supervisor: Asst. Prof. Dr. Abdullah ULAS
Co-Supervisor: Prof. Dr. Kahraman ALBAYRAK
April 2005, 164 pages
Development of the combustor of a ramjet can be achieved by connected pipe testing. Connected pipe testing is selected for combustor testing because pressure, temperature, Mach number, air mass flow rate can be simulated by this type of testing. Real time trajectory conditions and transition from rocket motor (booster) to ramjet operation can also be tested. The biggest advantage of connected pipe testing is the low operation cost and simplicity. Air mass flow rate requirement is less than the others which requires less air storage space and some components like supersonic nozzle and ejector system is not necessary.
In this thesis, design of a connected pipe test facility is implemented. Three main systems are analyzedair storage system, air heater system and test stand. Design of air storage system includes the design of pressure vessel and pressure &
flow regulation system. Pressure and flow regulation system is needed to obtain the actual flow properties that the combustor is exposed to during missile flight. Alternatives for pressure and air mass flow rate regulation are considered in this study. Air storage system designed in this thesis is 27.8 m3 at 50 bar which allows a test duration of 200 seconds at an average mass flow rate of 3 kg/s. Air heater system is utilized to heat the air to simulate the aerodynamic heating of the inlet. Several different combustion chamber configurations with different flame holding mechanisms are studied. The most efficient configuration is selected for this study. Combustion analysis of the air heater is performed by FLUENT CFD Code. Combustion process and air heater designs are validated using experimental data. Designed air heater system is capable of supplying air at a temperature range of 400-1000 K and mass flow rate range of 1.5-8 kg/s at Mach numbers between 0.1-0.5 and pressure between 2-8 bar. Finally the design of the test stand and ramjet combustor analysis are completed. 3D CAD models of the test stand are generated. Ramjet combustor that will be tested in the test setup is modeled and combustion analysis is performed by FLUENT CFD Code. The ramjet engine cruise altitude is 16 km and cruise Mach number is 3.5. Key-words: Air Breathing Engines, Ramjet, Connected Pipe, Direct Connect, Vitiator.
9
Piper, Ross H. "Design and testing of a combustor for a turbo-ramjet for UAV and missile applications." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2003. http://library.nps.navy.mil/uhtbin/hyperion-image/03Mar%5FPiper.pdf.
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Thesis (M.S. in Aeronautical Engineering)--Naval Postgraduate School, March 2003.Thesis advisor(s): Garth V. Hobson, Raymond P. Shreeve. Includes bibliographical references (p. 81-82). Also available online.
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Gabruk, Robert S. "The characterization of the flowfield of a dump combustor." Thesis, This resource online, 1990. http://scholar.lib.vt.edu/theses/available/etd-05092009-040628/.
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Ferguson, Kevin M. "Design and cold flow evaluation of a miniature Mach 4 Ramjet." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2003. http://library.nps.navy.mil/uhtbin/hyperion-image/03Jun%5FFerguson.pdf.
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Thesis (M.S. in Aeronauticl Engineering)--Naval Postgraduate School, June 2003.Thesis advisor(s): Garth V. Hobson, Raymond P. Shreeve. Includes bibliographical references (p. 67). Also available online.
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Redford, Tim. "Effects of incomplete fuel-air mixing on the performance characteristics of mixed compression, shock-induced combustion ramjet, shcramjet, engines." Thesis, National Library of Canada = Bibliothèque nationale du Canada, 1998. http://www.collectionscanada.ca/obj/s4/f2/dsk2/tape17/PQDD_0010/MQ34109.pdf.
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Baig, Saood Saeed. "A simple moving boundary technique and its application to supersonic inlet starting /." Thesis, McGill University, 2008. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=112555.
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In this thesis, a simple moving boundary technique has been suggested, implemented and verified. The technique may be considered as a generalization of the well-known "ghost" cell approach for boundary condition implementation. According to the proposed idea, the moving body does not appear on the computational grid and is allowed to move over the grid. The impermeable wall boundary condition is enforced by assigning proper gasdynamic values at the grid nodes located inside the moving body close to its boundaries (ghost nodes). The reflection principle taking into account the velocity of the boundaries assigns values at the ghost nodes. The new method does not impose any particular restrictions on the geometry, deformation and law of motion of the moving body.The developed technique is rather general and can be used with virtually any finite-volume or finite-difference scheme, since the modifications of the schemes themselves are not required. In the present study the proposed technique has been incorporated into a one-dimensional non-adaptive Euler code and a two-dimensional locally adaptive unstructured Euler code.
It is shown that the new approach is conservative with the order of approximation near the moving boundaries. To reduce the conservation error, it is beneficial to use the method in conjunction with local grid adaptation.
The technique is verified for a number of one and two dimensional test cases with analytical solutions. It is applied to the problem of supersonic inlet starting via variable geometry approach. At first, a classical starting technique of changing exit area by a moving wedge is numerically simulated. Then, the feasibility of some novel ideas such as a collapsing frontal body and "tractor-rocket" are explored.
14
Sykes, Harrison G. "COLD FLOW PERFORMANCE OF A RAMJET ENGINE." DigitalCommons@CalPoly, 2014. https://digitalcommons.calpoly.edu/theses/1334.
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The design process and construction of the initial modular ramjet attachment to the Cal Poly supersonic wind tunnel is presented. The design of a modular inlet, combustor, and nozzle are studied in depth with the intentions of testing in the modular ramjet. The efforts undertaken to characterize the Cal Poly supersonic wind tunnel and the individual component testing of this attachment are also discussed. The data gathered will be used as a base model for future expansion of the ramjet facility and eventual hot fire testing of the initial components. Modularity of the inlet, combustion chamber, and nozzle will allow for easier modification of the initial design and the designs ability to incorporate clear walls will allow for flow and combustion visualization once the performance of the hot flow ramjet is determined. The testing of the blank ramjet duct resulted in an error of less than 10% from predicted results. The duct was also tested with the modular inlet installed and resulted in between a 13-30% error based on the predicted results. Hot flow characteristics of the ramjet were not achieved, and the final cold flow test with the nozzle installed was a failure due to improper configuration of the nozzle. The errors associated with this testing can largely be placed on the poor performance of the Cal Poly supersonic wind tunnel and the alterations made to the testing in an attempt to accommodate these flaws. The final tests were halted for safety concerns and could continue after a thorough safety review.
15
Antoun, Sami Jamil. "Potential of the ramjet engine for hypersonic flight speeds /." The Ohio State University, 1993. http://rave.ohiolink.edu/etdc/view?acc_num=osu1487842372898003.
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Dupont, Benoît. "Conception du compresseur supersonique du Rim Rotor Rotary Ramjet Engine." Mémoire, Université de Sherbrooke, 2015. http://hdl.handle.net/11143/8823.
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La demande pour les ressources énergétiques est en hausse alors que leur disponibilité est en baisse. Dans ce contexte, l’industrie du transport et de l’énergie est à la recherche de petits moteurs efficaces et puissants et le Rim Rotor Rotary Ramjet Engine (R4E) pourrait correspondre à ces critères. Or, en ce moment, le potentiel de ce moteur est limité, car son compresseur supersonique entraîne des pertes d’efficacité lorsque le rotor tourne à son nombre de Mach tangentiel optimal qui est de 2.
Le présent mémoire compile toutes les notions requises pour comprendre le fonctionnement d’un compresseur supersonique lors de son démarrage et de concevoir le compresseur le plus approprié pour le R4E, tant en démarrage qu’en régime permanent. Pour se faire, des concepts de cascades inspirés des compresseurs et des méthodes de démarrage des moteurs ramjet actuels ont été générés et validés à l’aide de modèles analytiques. Les concepts sont par la suite essayés expérimentalement sous la forme de cascades à l’aide d’une soufflerie supersonique.
Bien que le modèle analytique montre que les cascades munies de canaux de purge soient plus performantes et plus robustes en conditions off-design, ces dernières n’ont jamais démarré lors des expérimentations même si les canaux ont été agrandis et multipliés. Ainsi, parmi tous les concepts essayés, celui qui démarre par survitesse et qui comporte des canaux de succion de couche limite à son col a donné les meilleurs résultats. Il est très stable et permet d’obtenir un ratio de pression statique de 4.25 et un recouvrement de pression totale de 89 %, pour une efficacité isentropique de 92 % à un nombre de Mach tangentiel de 2. Par contre, il est à noter qu’il n’a pas été possible de mesurer la pression totale. Elle a plutôt été estimée à partir des images de strioscopie tirées lors des essais.
Comme on ne dispose pas d’une structure permettant d’essayer le compresseur rotatif à Mach 2, il a fallu approximer l’influence de l’accélération centrifuge sur l’écoulement de la cascade et trouver un moyen d’intégrer le nouvel aubage à la roue. Un modèle permettant d’estimer les paramètres d’une couche limite se développant sur une plaque plane en rotation a permis de déduire que l’accélération transverse n’aurait qu’un effet légèrement favorable, puisqu’il permet d’amincir l’épaisseur de déplacement, réduisant ainsi les risques d’interaction en la couche limite et les chocs. Finalement, les canaux de succion de couche limite du compresseur pourraient permettre d’alimenter un système de refroidissement qui limiterait la température à la jante à 820 K.
Le R4E pourrait devenir l’avenir des systèmes de régénération électrique pour les véhicules hybrides. Il serait aussi intéressant pour une utilisation dans les petites centrales thermiques des régions éloignées. Ce grand potentiel d’utilisation provient de la grande densité de puissance du moteur, de sa simplicité et de son très faible coût de fabrication et de maintenance.
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Picard, Mathieu. "Dynamique des gaz et combustion du Rim-Rotor Rotary Ramjet Engine (R4E)." Mémoire, Université de Sherbrooke, 2011. http://savoirs.usherbrooke.ca/handle/11143/1607.
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Le Rim-Rotor Rotary Ramjet Engine (R4E) a le potentiel de remplacer les turbines à gaz de 1 MW et moins en offrant : (1) une densité de puissance de 7.6 kW/kg, soit le double des turbines à gaz actuelles, (2) une meilleure fiabilité et un moindre coût par sa pièce mobile unique et (3) une efficacité de plus de 20 %, soit similaire aux turbines à gaz de cette puissance.Le R4E convertit la grande vitesse tangentielle du mélange air-carburant, idéalement 1000 m/s, en une grande pression dans la chambre de combustion. La combustion des réactifs augmente le volume du gaz ce qui force les produits à sortir à une grande vitesse tangentielle. La poussée générée est récupérée en travail mécanique à l'arbre directement ou est convertie en électricité. Ce travail présente la conception de la géométrie des propulseurs à l'aide d'un modèle 1D généralisé basé sur l'analyse préliminaire, ainsi que la validation expérimentale d'un prototype faisant la preuve de concept du R4E à travers 5 étapes principales s'étendant sur 2 versions du prototype : (1) la friction aérodynamique, (2) l'écoulement dans le moteur, (3) l'allumage, (4) la combustion et (5) la démonstration de la puissance nette des ailettes. La friction aérodynamique de la paroi externe du Rim-Rotor dépasse de 35 % les modèles actuels ce qui en fait le mécanisme de perte le plus important.Le débit massique dans le moteur est de 30 % inférieur à la valeur estimée par le modèle 1D pour la géométrie testée. La puissance de trainée des statoréacteurs sans combustion mesurée est en ligne avec la puissance prédite pour un débit massique corrigé expérimentalement. En ce qui concerne la combustion dans le moteur, le champ centrifuge extrême domine le mécanisme de propagation de la flamme. Un modèle simple de flottaison est utilisé pour prédire la longueur du front de flamme, représentant les produits chauds qui ont tendance à"flotter" sur les réactifs froids. Un modèle numérique est élaboré pour valider la propagation de la flamme jusqu'à une accélération centrifuge de 1.1 million de g et montre une bonne corrélation avec le modèle simple. Une efficacité de combustion de 85% est démontrée avec un second prototype pour une accélération centrifuge jusqu'à 284 000 g, soit 25 fois supérieures à la plus grande valeur testée dans la littérature. Une fois la combustion stabilisée, ce prototype a été en mesure de produire une légère poussée, une première pour les moteurs à statoréacteur rotatif.
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Rancourt, David. "Analyse structurelle et validation expérimentale d'un Rim-Rotor Rotary Ramjet Engine (R4E)." Mémoire, Université de Sherbrooke, 2011. http://savoirs.usherbrooke.ca/handle/11143/1612.
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Le Rim-Rotor Rotary Ramjet Engine (R4E) est un moteur très haute densité de puissance utilisant des statoréacteurs en rotation pour produire un couple à grande vitesse angulaire.Le design structurel d'une première génération de R4E est présenté dans ce mémoire ainsi qu'une validation expérimentale avec combustion haute température. Ces travaux s'imbriquent dans un programme de recherche où l'objectif ultime est de démontrer expérimentalement qu'il est possible de produire de la puissance positive de ce type de moteur. La structure principale du moteur est basée sur l'utilisation d'un Rim-Rotor, un anneau de Carbone-PEEK unidirectionnel, qui reprend partiellement le chargement des propulseurs en compression. Un moyeu en aluminium en une pièce inclut les propulseurs et supporte le système d'allumage inductif intégré à la structure. Ce dernier a été caractérisé indépendamment afin de connaître l'effet des paramètres tels la distance entre les électrodes sur la puissance et l'énergie des étincelles.Le concept final proposé pèse 76 g, ne contient que 5 pièces dans un assemblage unique et peut résister à une vitesse tangentielle de 330 m/s (120 krpm) au niveau des propulseurs lors d'une combustion d'hydrogène de 1 sec. Un autre concept présenté est conçu pour résister 560 m/s (200 krpm) pour des durées de combustion très courtes, sans échauffement significatif des composants. Un modèle structurel analytique est proposé et validé par un modèle numérique ainsi que des essais expérimentaux sans combustion réalisés jusqu'à 188 krpm sans rupture.Le prototype conçu pour la combustion est validé par rapport à ses paramètres de conception et une rupture des pales de turbine survient tel que prédit par le modèle couplé thermique-structurel numérique. Les recherches ont démontré que le concept d'un R4E est viable et qu'il a le potentiel d'atteindre une vitesse tangentielle de près de 1000 m/s en utilisant des matériaux disponibles aujourd'hui. Les dissimilitudes d'expansion thermique entre les composantes, la différence de rigidité entre les pièces de l'assemblage ainsi que le transfert de chaleur vers le Rim-Rotor ont été identifiés comme des considérations importantes pour les futurs concepts de R4E.
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Krikellas, Dimitrios. "Improvement of the performance of a turbo-ramjet engine for UAV and missile applications." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2003. http://library.nps.navy.mil/uhtbin/hyperion-image/03Dec%5FKrikellas.pdf.
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Thesis (M.S. in Applied Physics and M.S. in Aeronautical Engineering)--Naval Postgraduate School, December 2003.Thesis advisor(s): Garth V. Hobson, Kai E. Woehler. Includes bibliographical references (p. 133). Also available online.
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Fruge, Keith J. "Design and testing of a caseless solid-fuel integral-rocket ramjet engine for use in small tactical missiles." Thesis, Monterey, California. Naval Postgraduate School, 1991. http://hdl.handle.net/10945/28115.
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Bond, Ryan Bomar. "Reynolds-Averaged Navier-Stokes Analysis of the Flow through a Model Rocket-Based Combined Cycle Engine with an Independently-Fueled Ramjet Stream." NCSU, 2003. http://www.lib.ncsu.edu/theses/available/etd-08132003-171258/.
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A new concept for the low speed propulsion mode in rocket based combined cycle (RBCC) engines has been developed as part of the NASA GTX program. This concept, called the independent ramjet stream (IRS) cycle, is a variation of the traditional ejector ramjet (ER) design and involves the injection of hydrogen fuel directly into the air stream, where it is ignited by the rocket plume. Experiments and computational fluid dynamics (CFD) are currently being used to evaluate the feasibility of the new design. In this work, a Navier-Stokes code valid for general reactive flows is applied to the model engine under cold flow, ejector ramjet, and IRS cycle operation. Pressure distributions corresponding to cold-flow and ejector ramjet operation are compared with experimental data. The engine response under independent ramjet stream cycle operation is examined for different reaction models and grid sizes. The engine response to variations in fuel injection is also examined. Mode transition simulations are also analyzed both with and without a nitrogen purge of the rocket. The solutions exhibit a high sensitivity to both grid resolution and reaction mechanism, but they do indicate that thermal throat ramjet operation is possible through the injection and burning of additional fuel into the air stream. The solutions also indicate that variations in fuel injection location can affect the position of the thermal throat. The numerical simulations predicted successful mode transition both with and without a nitrogen purge of the rocket; however, the reliability of the mode transition results cannot be established without experimental data to validate the reaction mechanism.
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O'Byrne, Sean Brendan. "Examination of transient mixing and combustion processes in a supersonic combustion ramjet engine." Master's thesis, 1997. http://hdl.handle.net/1885/145993.
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Devaraj, Manoj Kumar K. "Physical insights into unstart dynamics of a hypersonic mixed compression intake." Thesis, 2021. https://etd.iisc.ac.in/handle/2005/5655.
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Hypersonic air-breathing cruise vehicles powered by supersonic combustion ramjet engines
are the potential candidate for future space and defense applications. The air intake of the
scramjet engine is a vital component that uses shock waves to compress the air to pressure
and temperatures suitable for supersonic combustion. Understanding the unstart dynamics of
such intakes is of prime importance for the seamless operation of scramjet intakes. While the
unstart dynamics in supersonic intakes are studied widely by various researchers, only a few
such studies are reported in hypersonic intakes. The mechanisms associated with the same are
not clearly understood. In the current work, a design optimization framework is established
by coupling (a) oblique-shock theory and Non-dominated Sorting Genetic Algorithm II
(NSGA-II) and (b) Computational fluid dynamics (CFD) and NSGA - II to minimize total
pressure loss and maximize intake exit temperature of planar mixed compression intake at
a design Mach number of 6. The ramp and cowl angles constitute the design space. The
intake with maximum exit temperature is chosen to study its unstart dynamics using a
combination of experiments in a hypersonic wind tunnel (M = 6 and Re = 8.86 × 106/m)
and unsteady numerical investigations using the open-source suite SU2. The intake model
is equipped with a movable cowl and flap to study the internal contraction and throttling
induced unstart. Simultaneous pressure measurements and schlieren flow visualization are
carried out to study unsteady flow physics associated with intake unstart. The dynamic
content in the flow is analyzed using Fast Fourier Transform (FFT) and spectrogram of the
unsteady pressure signal and Dynamic Mode Decomposition (DMD) of the schlieren images
and density contours.
In this work, two different modes of shock oscillation during unstart are observed when
the flap is moved while the cowl is held stationary. At ICR = 1.19, the intake shows started
behavior for throttling ratio up to 0.31, and a dual behavior, where it remains started in
dynamic flap runs but unstarted in fixed flap runs for throttling ratios of 0.35 and 0.42. The
intake exhibits a staged evolution to a large amplitude oscillatory unstart for throttling ratios
of 0.55 and 0.69, with frequencies of 950 and 1100 Hz, respectively. A staged evolution (5
stages) to a subsonic spillage oscillatory unstart is detailed using corroborative evidence from
both time-resolved schlieren and pressure measurements. The ramp side separation bubble
drives the high amplitude oscillatory unstart. At ICR = 1.37, the shear layer emanating
from the triple point of shock interaction drives the low amplitude oscillatory unstart with a
dominant frequency of about 3.7 kHz for a throttling ratio of 0.69. A criterion for demarcating
the modes of unstart is evolved using current and previous data. The actual shock on lip
condition during started operation demarcates the two modes of oscillatory unstart. Unsteady
numerical computations are performed to study the effect of enthalpy on the unstart frequency.
The frequency of unstart varies linearly with stagnation acoustic speed and is an appropriate
velocity scale. During unstart, the extent of the subsonic region is the appropriate length
scale to be used in the quarter-wave resonance model to estimate unstart frequency pertaining
to high mechanical blockage
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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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Sivaprakasam, M. "Numerical Simulations Of Two-Phase Reacting Flow In A Cavity Combustor." Thesis, 2010. https://etd.iisc.ac.in/handle/2005/2011.
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In the present work, two phase reacting flow in a single cavity Trapped Vortex Combustor (TVC) is studied at atmospheric conditions. KIVA-3V, numerical program for simulating three dimensional compressible reacting flows with sprays using Lagrangian-Drop Eulerian-fluid procedure is used. The stochastic discrete droplet model is used for simulating the liquid spray. In each computational cell, it is assumed that the volume occupied by the liquid phase is very small. But this assumption of very low liquid volume fraction in a computational cell is violated in the region close to the injection nozzle. This introduces grid dependence in predictions of liquid phase in the region close to the nozzle in droplet collision algorithm, and in momentum coupling between the liquid and the gas phase. Improvements are identified to reduce grid dependence of these algorithms and corresponding changes are made in the standard KIVA-3V models.
Pressure swirl injector which produces hollow cone spray is used in the current study along with kerosene as the liquid fuel. Modifications needed for modelling pressure swirl atomiser are implemented. The Taylor Analogy Breakup (TAB) model, the standard model for predicting secondary breakup is improved with modifications required for low pressure injectors. The pressure swirl injector model along with the improvements is validated using experimental data for kerosene spray from the literature.
Simulations of two phase reacting flow in a single cavity TVC are performed and the temperature distribution within the combustor is studied. In order to identify an optimum configuration with liquid fuel combustion, the following parameters related to fuel and air such as cavity fuel injection location, cavity air injection location, Sauter Mean Diameter (SMD) of injected fuel droplets, velocity of the fuel injected are studied in detail in order to understand the effect of these parameters on combustion characteristics of a single cavity TVC.
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Sivaprakasam, M. "Numerical Simulations Of Two-Phase Reacting Flow In A Cavity Combustor." Thesis, 2010. http://etd.iisc.ernet.in/handle/2005/2011.
Full textAbstract:
In the present work, two phase reacting flow in a single cavity Trapped Vortex Combustor (TVC) is studied at atmospheric conditions. KIVA-3V, numerical program for simulating three dimensional compressible reacting flows with sprays using Lagrangian-Drop Eulerian-fluid procedure is used. The stochastic discrete droplet model is used for simulating the liquid spray. In each computational cell, it is assumed that the volume occupied by the liquid phase is very small. But this assumption of very low liquid volume fraction in a computational cell is violated in the region close to the injection nozzle. This introduces grid dependence in predictions of liquid phase in the region close to the nozzle in droplet collision algorithm, and in momentum coupling between the liquid and the gas phase. Improvements are identified to reduce grid dependence of these algorithms and corresponding changes are made in the standard KIVA-3V models.
Pressure swirl injector which produces hollow cone spray is used in the current study along with kerosene as the liquid fuel. Modifications needed for modelling pressure swirl atomiser are implemented. The Taylor Analogy Breakup (TAB) model, the standard model for predicting secondary breakup is improved with modifications required for low pressure injectors. The pressure swirl injector model along with the improvements is validated using experimental data for kerosene spray from the literature.
Simulations of two phase reacting flow in a single cavity TVC are performed and the temperature distribution within the combustor is studied. In order to identify an optimum configuration with liquid fuel combustion, the following parameters related to fuel and air such as cavity fuel injection location, cavity air injection location, Sauter Mean Diameter (SMD) of injected fuel droplets, velocity of the fuel injected are studied in detail in order to understand the effect of these parameters on combustion characteristics of a single cavity TVC.
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Cheng-YuChiang and 蔣政育. "Studies of Supersonic Ramjet Engine Cooling Technology by using Endothermic Hydrocarbon Fuels." Thesis, 2018. http://ndltd.ncl.edu.tw/handle/hr7e4b.
Full textAbstract:
碩士國立成功大學
航空太空工程學系
106
Scramjet engine thermal management system technology, which combined convection cooling by passing fuel through internal channel structures and the fuel endothermic cracking reaction, is one of the main efforts of our current national defense science and technology. Such thermal management system with convection cooling and fuel endothermic cracking reaction has been regarded as a key technology of scramjet engine around the world. With the implementation of this research, it is expected to develop an analysis tool of scramjet engine thermal management system that can help researchers in the field of scramjet engine in their design of thermal management systems. The research developed a high efficiency thermal management system that can be used for the next generation hypersonic speed vehicle scramjet's cooling problems. Scramjet engine thermal management systems analysis technology platform development in this research, in addition to the implementation of this research, it can also promote the possibility of technical autonomy. This research will analyze the thermal management system designs for the scramjet engine in hypersonic vehicle based on the heat flux needs and surface temperature limits. It is found that the closer working pressure of the hydrocarbon endothermic fuel in the cooling pipe is to the critical pressure, the more severe physical properties changes. In this process, the heat transfer deterioration caused by the decreased Cp and ρ. Various endothermic fuel thermal management system will be designed and evaluated to meet the heat transfer needs. Comparison with empirical heat transfer correlations, and then make new heat transfer correlation of JP-10 hydrocarbon fuel to facilitate Compared with the heat transfer empirical formula in the literature, the physical properties (density, specific heat, thermal conductivity, viscosity coefficient) and pressure parameters of the JP-10 hydrocarbon fuel can not be used in the JP-10 hydrocarbon fuel. Therefore, the heat transfer empirical formula will be made in this paper to facilitate the consideration of setting parameters.