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1
Hackaday, Gary L. "Thrust augmentation for a small turbojet engine." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 1999. http://handle.dtic.mil/100.2/ADA362981.
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Thesis (M.S. in Aeronautical Engineering) Naval Postgraduate School, March 1999.Thesis advisor(s): Garth V. Hobson. "March 1999". Includes bibliographical references (p. 75). Also available online.
2
Andreou, Loukas. "Performance of a ducted micro-turbojet engine." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 1999. http://handle.dtic.mil/100.2/ADA370851.
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Thesis (M.S. in Aeronautical Engineering) Naval Postgraduate School, September 1999."September 1999". Thesis advisor(s): Garth V. Hobson. Includes bibliographical references (p. 79). Also available online.
3
Fahlström, Simon, and Rikard Pihl-Roos. "Design and construction of a simple turbojet engine." Thesis, Uppsala universitet, Industriell teknik, 2016. http://urn.kb.se/resolve?urn=urn:nbn:se:uu:diva-303970.
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This project deals with researching, designing and building jet-engines. A simple turbojet engine was designed and construction was begun. The design was made by studying the work done by industry and researchers over the course of the history of jet engines. The methods were then discussed and chosen in a way that would simplify the design work as well as the construction of the engine. The goal was to create a self-sustaining combustion within the engine. The design settled upon consists of a radial compressor, an annular combustion chamber and an axial turbine. Since the compressor would have been the most difficult part to machine the decision was made early on to use the compressor from a turbocharger out of an automotive engine. Upon further study it was discovered that the characteristics of this compressor was not compatible with the rest of the design, as the compressor was made for an RPM range outside of what we could achieve and the compression ratio was too low. Most of the rest of the engine had already been built, and there was not enough time to design and build another compressor so work was aborted on the engine.
4
Yarlagadda, Santosh. "Performance Analysis of J85 Turbojet Engine Matching Thrust with Reduced Inlet Pressure to the Compressor." University of Toledo / OhioLINK, 2010. http://rave.ohiolink.edu/etdc/view?acc_num=toledo1271367584.
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Bell, Jabin Todd. "Measurements of forced and unforced aerodynamic disturbances in a turbojet engine." Thesis, Massachusetts Institute of Technology, 1993. http://hdl.handle.net/1721.1/46423.
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6
Ceylanoglu, Arda. "An Accelerated Aerodynamic Optimization Approach For A Small Turbojet Engine Centrifugal Compressor." Master's thesis, METU, 2009. http://etd.lib.metu.edu.tr/upload/12611371/index.pdf.
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Centrifugal compressors are widely used in propulsion technology. As an important part of turbo-engines, centrifugal compressors increase the pressure of the air and let the pressurized air flow into the combustion chamber. The developed pressure and the flow characteristics mainly affect the thrust generated by the engine.
The design of centrifugal compressors is a challenging and time consuming process including several tests, computational fluid dynamics (CFD) analyses and optimization studies. In this study, a methodology on the geometry optimization and CFD analyses of the centrifugal compressor of an existing small turbojet engine are introduced as increased pressure ratio being the objective. The purpose is to optimize the impeller geometry of a centrifugal compressor such that the pressure ratio at the maximum speed of the engine is maximized. The methodology introduced provides a guidance on the geometry optimization of centrifugal impellers supported with CFD analysis outputs.
The original geometry of the centrifugal compressor is obtained by means of optical scanning. Then, the parametric model of the 3-D geometry is created by using a CAD software. A design of experiments (DOE) procedure is applied through geometrical parameters in order to decrease the computation effort and guide through the optimization process. All the designs gathered through DOE study are modelled in the CAD software and meshed for CFD analyses. CFD analyses are carried out to investigate the resulting pressure ratio and flow characteristics.
The results of the CFD studies are used within the Artificial Neural Network methodology to create a fit between geometric parameters (inputs) and the pressure ratio (output). Then, the resulting fit is used in the optimization study and a centrifugal compressor with higher pressure ratio is obtained by following a single objective optimization process supported by design of experiments methodology.
7
Pavelec, Sterling Michael. "The development of turbojet aircraft in Germany, Britain, and the United States : a multi-national comparison of aeronautical engineering, 1935-1946 /." The Ohio State University, 2004. http://www.ohiolink.edu/etd/send-pdf.cgi/Pavelec%20Sterling%20Michael.pdf?acc_num=osu1082396007.
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Polat, Cuma. "An Electronic Control Unit Design For A Miniature Jet Engine." Master's thesis, METU, 2009. http://etd.lib.metu.edu.tr/upload/12611442/index.pdf.
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Gas turbines are widely used as power sources in many industrial and transportation applications. This kind of engine is the most preferred prime movers in aircrafts, power plants and some marine vehicles. They have different configurations according to their mechanical constructions such as turbo-prop, turbo-shaft, turbojet, etc. These engines have different efficiencies and specifications and some advantages and disadvantages compared to Otto-Cycle engines. In this thesis, a small turbojet engine is investigated in order to find different control algorithms.
AMT Olympus HP small turbojet engine has been used to determine the mathematical model of a gas turbine engine. Some important experimental data were taken from AMT Olympus engine by making many experiments. All components of the engine have been modeled by using laws of thermodynamics and some arithmetic calculations such as numerical solution of nonlinear differential equations, digitizing compressor and turbine map etc. This mathematical model is employed to create control algorithm of the engine. At first, standard control strategies had been considered such as P (proportional), PI (proportional integral), and PID (proportional-integral-differential) controllers. Because of the nonlinearities in gas turbines, standard control algorithms are not commonly used in literature. At the second stage fuzzy logic controllers have been designed to control the engine efficiently. This control algorithm was combined with mathematical of the engine in MATLAB environment and input-output relations were investigated. Finally, fuzzy logic control algorithm was embedded into an electronic controller.
9
KAMARAJ, JAYACHANDRAN. "MODELING AND SIMULATION OF SINGLE SPOOL JET ENGINE." University of Cincinnati / OhioLINK, 2004. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1073935505.
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Avrat, Jan. "Metody údržby a diagnostiky lopatkových motorů." Master's thesis, Vysoké učení technické v Brně. Fakulta strojního inženýrství, 2009. http://www.nusl.cz/ntk/nusl-228602.
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This master`s thesis deals with methods of turbojet engines maintenance and diagnostics. First specify and review methods. Then choose acceptable and perspective methods of non-destructive testing. Based on this choosing will be designed modern service center.
11
Benyo, Theresa L. "Analytical and Computational Investigations of a Magnetohydrodynamic (MHD) Energy-Bypass System for Supersonic Turbojet Engines to Enable Hypersonic Flight." Kent State University / OhioLINK, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=kent1369153719.
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Cevik, Mert. "Desifn And Optimization Of A Mixed Flow Compressor Impeller Using Robust Design Methods." Master's thesis, METU, 2009. http://etd.lib.metu.edu.tr/upload/12611105/index.pdf.
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This is a study that is focused on developing an individual design methodology for a centrifugal impeller and generating a mixed flow impeller for a small turbojet engine by using this methodology. The structure of the methodology is based on the design, modeling and the optimization processes, which are operated sequentially. The design process consists of
engine design and compressor design codes operated together with a commercial design code. Design of Experiment methods and an in-house Neural Network code is used for the modeling phase. The optimization is based on an in-house code which is generated based on multidirectional search algorithm. The optimization problem is constructed by using the inhouse parametric design codes of the engine and the compressor. The goal of the optimization problem is to reach an optimum design which gives the best possible combination of the thrust and the fuel consumption for a small turbojet engine.
The final combination of the design parameters obtained from the optimization study are used in order to generate the final design with the commercial design code. On the last part of the thesis a comparison of the final design and a standard radial flow impeller is made in order to clarify the benefit of the study. The results have been showed that a mixed flow
compressor design is superior to a standard radial flow compressor in a small turbojet application.
13
Mutschall, Marcel. "Die Genauigkeit einer vereinfachten Berechnung der Steigzeit von Flugzeugen." Aircraft Design and Systems Group (AERO), Department of Automotive and Aeronautical Engineering, Hamburg University of Applied Sciences, 2018. http://d-nb.info/1175497711.
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Ziel - Die Zeit die ein Flugzeug benötigt, um auf eine bestimmte Höhe zu steigen (die
Steigzeit) kann mit einer Formel berechnet werden, die vereinfachend annimmt, dass die
Steiggeschwindigkeit über dem gesamten Steigflug mit zunehmender Höhe linear abnimmt.
Ziel der Untersuchung ist, zu ermitteln, ob die Annahme einer linear abnehmenden
Steiggeschwindigkeit realistisch ist bzw. welche Fehler sich aus der Annahme ergeben.
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Methode - Mit der Höhe ändern sich Parameter wie Luftdichte, Widerstand, Schub und damit
auch die optimale Fluggeschwindigkeit für den Steigflug. Die Parameter beeinflussen sich
dabei gegenseitig. Der Schub wird dabei nach drei unterschiedlichen Methoden berechnet,
gegeben von Bräunling, Scholz und Howe. Analysiert wird der Verlauf des Schubes mit der
Höhe und der Verlauf der Steiggeschwindigkeit mit der Höhe für jede der drei
Schubberechnungen. Abschließend wird für jede Schubberechnung die Steigzeit verglichen
wie sie sich ergibt a) aus der einfachen Formel und b) aus einer Integrationsberechnung, bei
der der Verlauf der Steiggeschwindigkeit durch eine Funktion beschrieben wird.
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Ergebnisse - Die drei Schubberechnungen liefern ausgehend vom gleichen Startschub
unterschiedliche Schübe in der Höhe. In die Methode nach Bräunling gehen mehr Parameter
ein als in die anderen beiden Methoden. Es kann angenommen werden, dass die Methode
nach Bräunling genauer ist, der Beweis kann aber nicht geführt werden. Der Schub nach
Scholz und Howe fällt nahezu linear mit der Höhe ab. Der Schubverlauf nach Bräunling zeigt
eine deutliche Nichtlinearität. Es wird die Steigzeit von 0 km auf 11 km Höhe berechnet nach
a) und b), mit jeder der drei Schubberechnungen. Dabei wird jeweils der Unterschied in der
Steigzeit ermittelt. Aufgrund der Nichtlinearität im Schubverlauf zeigt die Methode nach
Bräunling dann auch den größten Unterschied zwischen den Berechnungsmethoden von
7,1 %. Bei einer Schubberechnung nach Scholz ergeben sich 1,7 % und nach Howe 1,4 %.
Wenn bereits zu Beginn Vereinfachungen, z.B. bezüglich des Triebwerksschubes,
vorgenommen wurden, ist es in Hinblick auf den Aufwand und die zu erreicheneden
Ergebnisse möglich, und zum Teil sinnvoll, die Berechnungen der Steigzeit mittels linearer
Abnahme der vertikalen Geschwindigkeit durchzuführen. Es wird ausdrücklich darauf
hingewiesen, dass es hier um den Vergleich von zwei Methoden zur Berechnung der Steigzeit
geht und nicht um die Bewertung von Methoden zur Schubberechnung (für die keine
Vergleichswerte vorlagen).
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Praktischer Nutzen - Es konnte festgestellt werden, dass eine einfache Formel zur
Berechnung der Steigzeit mit geringem Fehler angewandt werden kann - insbesondere wenn
Methoden zur Schubberechnung vorliegen, bei denen der Schub annähernd linear mit der
Höhe abnimmt. Bei großem Aufwand und realitätsnaher Betrachtung, z.B. nach Bräunling,
führt der lineare Ansatz jedoch zu einem zu großen Fehler. Hierfür sollte die Berechnung der
Steigzeit mittels Integration durchgeführt werden.
14
Rivera, Gilbert D. "Turbochargers to small turbojet engines for uninhabited aerial vehicles." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 1998. http://handle.dtic.mil/100.2/ADA346353.
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Thesis (Degree of Aeronautical and Astronautical Engineer) Naval Postgraduate School, June 1998.Thesis advisor(s): Garth V. Hobson, David W. Netzer. "June 1998." Includes bibliographical references (p. 73). Also available online.
15
Natter, Bernard. "Etude de l'usure et de l'ecaillage de pieces d'un reacteur d'avion par la technique d'activation en couches superficielles." Université Louis Pasteur (Strasbourg) (1971-2008), 1987. http://www.theses.fr/1987STR13223.
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Utilisation de la technique d'activation en couches superficielles. Formation du radioelement cobalt 56 par reaction nucleaire. Etude du profil de concentration et mesure in situ d'usure et d'ecaillage (cas des pompes a carburant). Aspects metallurgiques. Application a l'industrie aeronautique
16
Bensel, Artur. "Characteristics of the Specific Fuel Consumption for Jet Engines." Aircraft Design and Systems Group (AERO), Department of Automotive and Aeronautical Engineering, Hamburg University of Applied Sciences, 2018. http://d-nb.info/1175791237.
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Purpose of this project is a) the evaluation of the Thrust Specific Fuel Consumption (TSFC) of jet engines in cruise as a function of flight altitude, speed and thrust and b) the determination of the optimum cruise speed for maximum range of jet airplanes based on TSFC characteristics from a). Related to a) a literature review shows different models for the influence of altitude and speed on TSFC. A simple model describing the influence of thrust on TSFC seems not to exist in the literature. Here, openly available data was collected and evaluated. TSFC versus thrust is described by the so-called bucket curve with lowest TSFC at the bucket point at a certain thrust setting. A new simple equation was devised approximating the influence of thrust on TSFC. It was found that the influence of thrust as well as of altitude on TSFC is small and can be neglected in cruise conditions in many cases. However, TSFC is roughly a linear function of speed. This follows already from first principles. Related to b) it was found that the academically taught optimum flight speed (1.316 times minimum drag speed) for maximum range of jet airplanes is inaccurate, because the derivation is based on the unrealistic assumption of TSFC being constant with speed. Taking account of the influence of speed on TSFC and on drag, the optimum flight speed is only about 1.05 to 1.11 the minimum drag speed depending on aircraft weight. The amount of actual engine data was extremely limited in this project and the results will, therefore, only be as accurate as the input data. Results may only have a limited universal validity, because only four jet engine types were analyzed. One of the project's original value is the new simple polynomial function to estimate variations in TSFC from variations in thrust while maintaining constant speed and altitude.
17
Kranich, Niklas Brüge Felix. "Wartungskosten von Passagierflugzeugen bei verschiedener Triebwerksanzahl berechnet nach DOC-Methoden." Aircraft Design and Systems Group (AERO), Department of Automotive and Aeronautical Engineering, Hamburg University of Applied Sciences, 2018. http://d-nb.info/1175283754.
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Diese Projektarbeit versucht zu erklären,
warum Flugzeuge mit drei oder vier Triebwerken kaum noch verkauft werden.
Dabei wird insbesondere der Vermutung nachgegangen, dass Flugzeuge mit einer größeren Anzahl an Triebwerken höhere Wartungskosten haben könnten.
Zur Beantwortung der Frage werden sechs verschiedene Methoden zur Berechnung von Betriebskosten (Direct Operating Costs, DOC)
von Passagierflugzeugen herangezogen, die u.a. auch die Kosten der Triebwerkswartung abschätzen.
Vier dieser DOC-Methoden sind von Organisationen: Air Transport Association of America (ATA 1967), Deutsche Lufthansa (DLH 1982),
Association of European Airlines (AEA 1989), Airbus Industrie (AI 1989).
Zwei DOC-Methoden wurden an Universitäten entwickelt und sind von Jenkinson bzw. von Thorbeck (TU Berlin, TUB).
Weiterhin werden grundsätzliche flugmechanische Überlegungen angestellt und die Literatur durchgesehen,
die aber nur wenige Hinweise zur Beantwortung der Fragestellung enthält.
Die Gleichungen zur Berechnung der Triebwerkswartungskosten aller sechs Methoden werden dargelegt und erklärt.
Die Methoden unterscheiden sich stark in ihrer Komplexität.
Da die Methoden sich auf unterschiedliche Jahre beziehen werden die Kosten mit einem Inflationsfaktor auf das Jahr 2017 umgerechnet
und somit vergleichbar gemacht. Zum Vergleich werden weiterhin die Gleichungen zur Berechnung der Wartungskosten der Flugzeugzelle angegeben.
Zur Berechnung der Triebwerkswartungskosten wurden vier in der Größe vergleichbare Mittelstreckenflugzeuge ausgewählt:
B737-800, A318 (zwei Triebwerke), Jak-42 (drei Triebwerke), BAE 146-300 (vier Triebwerke).
Weiterhin wurden vier in der Größe vergleichbare Langstreckenflugzeug ausgewählt:
A330-300 (zwei Triebwerke), MD11-ER, TriStar (drei Triebwerke), A340-300 (vier Triebwerke).
Zum Vergleich eignen sich besonders der A330 und der A340 da die Technik, das Alter und die Abmaße sehr eng bei einander liegen.
Im Ergebnis wurde festgestellt, dass sich die Aufteilung der Wartungskosten zwischen Zelle und Triebwerken uneinheitlich zeigt.
Die AI-Methode ergibt im Vergleich viel zu hohe Triebwerkskosten.
Der Grund dafür ist die direkte Multiplikation von Schub mit den Lohnkosten.
Die AI-Methode muss daher bei der Endanalyse unberücksichtigt bleiben.
Bei den Mittelstreckenflugzeugen lieferten die Methoden nach AEA, DLH und TUB ähnliche Ergebnisse.
Bei den Langstreckenflugzeugen lieferten die AEA-Methode, DLH-Methode und die Methode nach Jenkison ähnliche Ergebnisse.
Empfohlen werden kann damit eine Berechnung mit der AEA-Methode, die auch öffentlich ist.
Für einen Endvergleich wurden für die Mittel- bzw. Langstrecke zu jeder Triebwerksanzahl nur jeweils ein Flugzeug einbezogen.
Mit dieser bereinigten Auswahl bei Flugzeugen und Methoden ergab sich für die Mittelstrecke eine leichte Abnahme der Triebwerkswartungskosten
mit der Triebwerksanzahl von nur 6,1 US$ pro Flugstunde pro Triebwerk (Zunahme von -6,1 US$/FH/Triebwerk).
Für die Langstrecke ergab sich eine leichte Zunahme der Triebwerkswartungskosten mit der Triebwerksanzahl von nur 32,5 US$ pro Flugstunde pro Triebwerk.
Damit konnte die eingangs genannte Vermutung über eine Zunahme der Triebwerkswartungskosten mit der Anzahl der Triebwerke nur zum Teil bestätigt werden.
Die Analyse zeigte, dass die Triebwerkswartungskosten von vielen Parametern abhängen, die Triebwerksanzahl ist nur ein Parameter von vielen.
Selbst ähnliche Flugzeuge liefern bei gleicher Triebwerkszahl daher Triebwerkswartungskosten,
die sich stark unterscheiden und die Abhängigkeit von der Triebwerkszahl wenig sichtbar werden lassen.
Es werden Vorschläge gemacht, welche anderen methodischen Ansätze hier Abhilfe schaffen könnten.
18
Ajrouche, Hussein. "Mesures optiques d'imbrûlés - applications aux émissions des moteurs Diesel et des réacteurs." Rouen, 1995. http://www.theses.fr/1995ROUES021.
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Une méthode optique in situ et non intrusive fondée sur la mesure simultanée de l'extinction et de la diffusion de la lumière par un nuage de particules a été mise au point. Elle a été utilisée pour caractériser le diamètre et la densité particulaire des suies émises par les véhicules Diesel et les turboréacteurs. Nous avons en particulier pu suivre dans un premier temps l'évolution des particules lors des accélérations successives d'un moteur Diesel. Il apparaît que le diamètre croît au cours de l'accélération alors que le nombre de particules tend à décroître. On retiendra comme ordre de grandeur typiques 50 nm et 510. 9 cm cubes à l'intérieur du tunnel. Ces résultats en régime instationnaire, que les méthodes classiques de filtration ne peuvent atteindre, sont à notre connaissance uniques. La première partie du travail a été réalisée au centre technique Citroën, le montage optique était directement fixé au tunnel de dilution recueillant les fumées émises par le pot d'échappement. La deuxième partie du travail décrit une cellule dédiée à la mesure des très faibles quantités d'imbrûlés, ici appliquée aux émissions des turboréacteurs. Compacte, elle utilise deux miroirs de renvoi pour générer un trajet optique total de 150 cm. Dans ces conditions des fractions volumiques inférieures a 10 puissance -8 ont été mesurées et leur évolution suivie en fonction de régime moteur. La discussion met en lumière l'influence des hypothèses faites lors du traitement des données, en particulier celles portant sur l'hypothèse de polydispersion et sur la valeur de l'indice complexe de réfraction des suies. Il est montré que l'usage des différentes valeurs de la littérature peut conduire par exemple à des variations de fraction volumique supérieures à 30%. Il est donc suggéré que l'accent soit mis sur la réduction de ces incertitudes. Sous ces conditions les méthodes optiques dont nous avons montré la faisabilité, y compris sur site industriel, seront un atout précieux dans les études touchant aux émissions d'aérosols et à leur suivi
19
Kwan, Pok Wang. "Flow management in heat exchanger installations for intercooled turbofan engines." Thesis, University of Oxford, 2011. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.711622.
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Gaeta, Richard Joseph Jr. "Liner impedance modification by varying perforate orifice geometry." Diss., Georgia Institute of Technology, 1998. http://hdl.handle.net/1853/12258.
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Nygaard, James Robert. "Understanding the behaviour of aircraft bearing steels under rolling contact loading." Thesis, University of Cambridge, 2015. https://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.709284.
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22
Giffard, Hermione S. "The development and production of turbojet aero-engines in Britain, Germany and the United States, 1936-1945." Thesis, Imperial College London, 2011. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.536008.
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Chesnakas, Christopher J. "Experimental studies in a supersonic through-flow fan blade cascade." Diss., Virginia Tech, 1991. http://hdl.handle.net/10919/39790.
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An investigation has been performed of the flow in a supersonic through-flow fan blade
cascade. The blade shapes are those of the baseline supersonic through-flow fan (STFF) under
investigation at the NASA Lewis Research Center. Measurements were made at an inlet Mach
number of 2.36 over a 15° range of incidence. Flowfield wave patterns were recorded using spark
shadowgraph photography and steady-state instrumentation was used to measure blade surface
pressure distributions and downstream total and static pressure distributions. A two-dimensional
LDV system was used to map the downstream flowfield. From these measurements, the integrated
loss coefficients are presented as a function of incidence angle along with analysis indicating the
source of losses in the STFF cascade. The results are compared with calculations made using a
two-dimensional, cell-centered, finite-volume, Navier-Stokes code with upwind options. Good
general agreement is found at design conditions, with lesser agreement at off-design conditions.
Analysis of the leading edge shock shows that the leading edge radius is a major source of losses
in STFF blades. Losses from the leading edge bluntness are convected downstream into the blade
wake, and are difficult to distinguish from viscous losses. Shock losses are estimated to account
for 70% to 80% of the losses in the STFF cascade.Ph. D.
24
Rajakuperan, E. "Experimental And Computational Investigations Of Underexpanded Jets From Elliptical Sonic Nozzles." Thesis, Indian Institute of Science, 1994. http://hdl.handle.net/2005/158.
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Three dimensional nozzles and jet flows have attracted the attention of many researchers due to their potential application to many practical devices. Rectangular nozzles are considered for short/vertical take off and landing aircrafts for achieving powered lift. Axisymmetric nozzles with lobes, tabs or slots and elliptical nozzles are considered for noise reduction in aircrafts and mixing augmentation in airbreathing rockets. Interaction of supersonic jets with solid
surface, as in the case of retro and ullage rockets in launch vehicles and interaction of multiple jets as in the case of launch vehicles with multiple booster rockets/multiple nozzle engines are of practical importance. Design of rockets and aircrafts employing these nozzles needs the understanding of the structure and behaviour of the complex three dimensional supersonic jets issuing from these nozzles. The problem is so complex that different investigators have addressed only some specific aspects of the problem and there is much more to be done to fully understand these flows. For example, in the case of rectangular nozzle with semi circular ends (known as elliptical nozzle), the investigations have been limited to a single nozzle of aspect ratio 3,0 and pressure ratio (ratio of the total pressure to ambient pressure) 3.0. Further, the measurements were made in the far field subsonic region beyond a distance of 20 times the equivalent nozzle radius (RJ.
For the present study, the elliptical sonic nozzle of the type mentioned above was chosen, as it offered simplicity for manufacturing and carrying out computations, but has all the complex features associated with the three dimensional jets. A systematic study to understand the mean flow structure and the effect of important governing parameters like
ratio and pressure ratio on the flow development process of the jet issuing from Navier-Stokes equations.
The experimental study revealed many interesting flow features. It was found that the Underexpanded jet issuing from elliptical sonic nozzle spreads rapidly in the minor axis plane while it maintains almost constant width or contracts in the major axis plane. However, the gross spreading of this jet is much higher compared to the axisymmetric jet. The higher spreading rates experienced in the minor axis plane compared to the major axis plane of this 'et, results in the jet width in the minor axis plane to become higher than that in the major axis plane. The longitudinal location, where this occurs is called the axis switching location. This kind of axis switching phenomenon is known to exist for subsonic elliptical jets. However, for the present supersonic jets, the axis switching locations are much closer to the nozzle exit compared to the subsonic cases reported. It was further found that this location strongly depends on the pressure and aspect ratios. A critical pressure ratio was found to exist for each nozzle at which the axis switching location is the farthest. Above the critical pressure ratio, the axis switching location was observed to move upstream with the increase in the pressure ratio and is controlled by the
complex interactions of shock and expansion waves near the nozzle exit. Below the critical pressure ratio, the axis switching location moves upstream with the decrease in pressure ratio and is controlled by some kind of instability in the minor axis plane.
The shock structure present in the underexpanded jet from an elliptical nozzle was also observed to depend on both pressure and aspect ratios. For some aspect ratios and pressure ratios, the shock pattern observed in both the major and minor axis planes are similar to that of an axisymmetric jet, where the incident barrel shock and the Mach reflection (from the edges of the Mach disk) are present. But for all other cases, this shock
continues to be seen only in the major axis plane. Whereas, in the minor axis plane, the incident shock is absent in the shock pattern.
Detailed measurement in the jet cross sectional planes, for the case of aspect ratio 2.0 nozzle, shows that the cross sectional shape changes along the length and it becomes almost a circle at the axis switching location. Further downstream, the jet spreads rapidly in the minor axis plane whereas no significant change in the width of the jet in the major axis plane is observed. Far downstream, the jet boundary appears like a distorted ellipse with its major and minor dimensions lying respectively in the minor and major axis planes of the nozzle. The elongated shape of the jet cross sections at locations downstream of the axis switching point gives the impression that the entire flow in the major axis plane is turned towards the minor axis plane. This effect appears to be predominant at high pressure ratios.
The computed near field shock structure in the planes of symmetry, pitot pressure distributions, cross sectional shape of the jet and the spreading pattern agree very well with the experimental results. In addition to this, the present computational method gives the detailed near field flow structure including the azimuthal extent of the incident shock, cross flow details and distributions of flow variables. It is shown that the present inviscid methodology can also predict the axis switching point accurately if it occurs before the formation of the Mach disk and it demonstrates that the jet growth phenomenon in the near field, atleast, is mainly controlled by the inviscid flow process. The computed results have shown that changes in the jet cross sectional shape in the near field is caused mainly by the interaction of compression and expansion waves with each other and with the constant pressure boundary. The inviscid method seems to be able to capture the complicated secondary cross flow structure (indicating presence of longitudinal vortices) of the elliptical jet.
The complex mean flow structure in the near field region of the jet issuing from elliptical nozzles and the effect of nozzle aspect ratio and pressure ratio on the structure are brought out clearly in the present study. The mechanism governing the spreading and the axis switching characteristics are also brought out. Thus the present experimental and computational investigations give a comprehensive understanding of the mean flow structure of the underexpanded jets issuing from elliptical nozzles. Further studies are required to understand the other aspects of the elliptical jets as well as other three-dimensional jets. Some of these studies are identified for future work.
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Hmad, Ouadie. "Evaluation et optimisation des performances de fonctions pour la surveillance de turboréacteurs." Thesis, Troyes, 2013. http://www.theses.fr/2013TROY0029.
Full textAbstract:
Cette thèse concerne les systèmes de surveillance des turboréacteurs. Le développement de tels systèmes nécessite une phase d’évaluation et d’optimisation des performances, préalablement à la mise en exploitation. Le travail a porté sur cette phase, et plus précisément sur les performances des fonctions de détection et de pronostic de deux systèmes. Des indicateurs de performances associés à chacune de ces fonctions ainsi que leur estimation ont été définis. Les systèmes surveillés sont d’une part la séquence de démarrage pour la fonction de détection et d’autre part la consommation d’huile pour la fonction de pronostic. Les données utilisées venant de vols en exploitation sans dégradations, des simulations ont été nécessaires pour l’évaluation des performances. L’optimisation des performances de détection a été obtenue par réglage du seuil sur la statistique de décision en tenant compte des exigences des compagnies aériennes exprimées en termes de taux de bonne détection et de taux d’alarme fausse. Deux approches ont été considérées et leurs performances ont été comparées pour leurs meilleures configurations. Les performances de pronostic de surconsommations d’huile, simulées à l’aide de processus Gamma, ont été évaluées en fonction de la pertinence de la décision de maintenance induite par le pronostic. Cette thèse a permis de quantifier et d’améliorer les performances des fonctions considérées pour répondre aux exigences. D’autres améliorations possibles sont proposées comme perspectives pour conclure ce mémoireThis thesis deals with monitoring systems of turbojet engines. The development of such systems requires a performance evaluation and optimization phase prior to their introduction in operation. The work has been focused on this phase, and more specifically on the performance of the detection and the prognostic functions of two systems. Performances metrics related to each of these functions as well as their estimate have been defined. The monitored systems are, on the one hand, the start sequence for the detection function and on the other hand, the oil consumption for the prognostic function. The used data come from flights in operation without degradation, simulations of degradation were necessary for the performance assessment. Optimization of detection performance was obtained by tuning a threshold on the decision statistics taking into account the airlines requirements in terms of good detection rate and false alarm rate. Two approaches have been considered and their performances have been compared for their best configurations. Prognostic performances of over oil consumption, simulated using Gamma processes, have been assessed on the basis of the relevance of maintenance decision induced by the prognostic. This thesis has allowed quantifying and improving the performance of the two considered functions to meet the airlines requirements. Other possible improvements are proposed as prospects to conclude this thesis
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Huang, Yi-Min, and 黃一民. "Study of a small Turbojet Engine Performance." Thesis, 2002. http://ndltd.ncl.edu.tw/handle/42250407371027985674.
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Hung, Lin Nan, and 林南宏. "Preliminary Design and Performance Test of cted Fan-Sprinkler Engine and Turbojet-Sprinkler Engine." Thesis, 1994. http://ndltd.ncl.edu.tw/handle/36382903996934168563.
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林家榮. "Combustion System Design and Testing for a Micro-Turbojet Engine." Thesis, 2012. http://ndltd.ncl.edu.tw/handle/27156629956737637959.
Full textAbstract:
碩士國立清華大學
動力機械工程學系
100
This study starts from the design theory of the holes on the combustion chamber liners . I designed the combustion chamber to reach the standards of KingTech K170-F micro-turbojet engines. The designed chamber was then assembled onto the K170-F turbo-engine in order to test the performance and analyze the combustion in the chamber. This study had three stages: (1) The performance of the K170-F turbo-engine was measured, and the turbine analysis program was used to calculate the conditions inside the turbo-engine, which in later stages would be referred to as the basics and restrictions; (2) Those conditions measured in last stage and the related theories were used to design a combustion chamber, in which details were all covered, including the geometrical chamber size, the positions and the diameters of the chamber liner holes, and the air flow rates of all combustion areas; (3) The chamber was assembled onto the K170-F turbo-engine, the performance of the running turbo-engine was measured, and then the testing results of the turbo-engine was compared with the designed chamber with the original ones in order to further analyze the combustion inside the chamber. The experimental results show that the chamber I designed can be applied to real turbo-engines and that the performing difference between the designed chamber and the original one is not significant. This study serves as a successful attempt to design a combustion chamber for a micro turbo-engine.
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Yang, Chih-Cheng, and 楊智丞. "Manufacture and Testing System Establishment of Hundred-Pound Micro Turbojet Engine." Thesis, 2018. http://ndltd.ncl.edu.tw/handle/mwpusx.
Full textAbstract:
碩士淡江大學
航空太空工程學系碩士班
106
This reseach investigated preliminary theoretical calculations, manufacturing process and control system design for a micro turbojet engine. Micro turbojet engine divided into inlet, compressor, diffuser, combustor chamber, turbine inlet guide vanes, turbine and nozzle. For theoretical calculations, we use one-dimensional ideal flow as the basis for calculation to decide aerodynamic parameter and aerodynamic shape by assuming pressure ratio of compressor and environment condition. The engine manufacturing has always been a difficult problem. We use turning, milling, casting, welding, deep hole drilling, electrical discharge machining, sheet metal methods to complete micro turbojet engine. In control system design, we write LabVIEW program with hardware by National Instrument and micro turbojet engine to complete process planning, and we also write a testing program to ensure whether control system operating normally.
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傅彥菳. "An investigation of compressor performance improvement for a small turbojet engine." Thesis, 2003. http://ndltd.ncl.edu.tw/handle/31311374972640649254.
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Tsai, Ming-Dar, and 蔡明達. "Study on Combustor Component Test and Diagnostics of a Micro Turbojet Engine." Thesis, 1998. http://ndltd.ncl.edu.tw/handle/97756838275258119851.
Full textAbstract:
碩士國立成功大學
航空太空工程學系
86
The combustor is an important component in turbojet engines. Gas turbinecombustor needs to be designed to meet many design requirements, includinghigh combustion efficiency over a wide operating envelop, stable operation,low pressure loss, low temperature pattern factor and low pollution emissions. In this study, a 60 lbf rated turbine combustor,P-60, is used to inves- tigate it''s possible reason to hot spot occurred in 45 degree to 135 de-gree at looking forward during the engine test. The diagnostics is sep-arated into cold flow test and hot flow test. In cold flow test, differen-tial pressures as airflow penetrating the liner hole are measured to far-ther understand its flow distribution. It is found that the jet dynamic pressure ratio K is found different from the large turbine combustors andK''s value seems too small that the discharge coefficient Cd is quite sen-sitive to flow condition which result in the variation of Cd and the flow distribution. A fuel atomizer performance is also evaluated on a spray test stand by both flow visualization technique and laser diffraction par-ticle analyzer. The spray characteristics are considerable satisfaction inthe design point operation.The pattern factor is about 1~2,and the combustion efficiency is above 90%after delicate calculation of airflow rate
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Wu, Yong Xian, and 吳永賢. "Design of a full operating envelope controller for a single-spool turbojet engine." Thesis, 1995. http://ndltd.ncl.edu.tw/handle/99399067384770772141.
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Ju, Cheng-Pin, and 茹承斌. "A Modeling Study for the Performance of a Single-Spooled Turbojet Engine Using Aerothermodynamics." Thesis, 1998. http://ndltd.ncl.edu.tw/handle/08030289051423881844.
Full textAbstract:
碩士中正理工學院
兵器工程研究所
86
ABSTRACT As the first step of a long-term serial study which has the ultimate goal of developing a simple and friendly simulation program for computing the integral performance of turbo-engines, this thesis presents some early stage efforts for the case of a single-spooled turbojet in order not to get lost in the complicated engine mechanisms. The major objective here is to work out a possible methodology on the basis of aerothermodynamic principles and partial data from real flight tests, so that it can be used to trace out the possible operating cycle of the turbojet. In order to justify the method validity, a so-called GasTurb software was adopted to generate data for comparisons and parameter plots for interpretations of the governing principles and relations followed. To accomplish with this latter part, performance variations of a real turbojet of the 3500 thrust level due to the changes of operating conditions was studied, with the expectation of being familiar with the analysis and design works of such engines. As in contents, the thesis begins with a detail analysis of the working principles for the engine components, and followed by an organization of these principles to yield a possible mathematical model for simulations, and then a case study of a chosen turbojet, and ended with some primary conclusions from the study.
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Tang-YuLai and 賴唐鈺. "A Study of Component Characteristics Analysis and Equilibrium Operating Prediction of a Turbojet Engine." Thesis, 2013. http://ndltd.ncl.edu.tw/handle/67335134585623468508.
Full textAbstract:
碩士國立成功大學
航空太空工程學系專班
101
The aircraft ordinary propulsion system is turbojet engine, which consists of compressor, combustor, turbine and exhaust nozzle. When engine is in stably working situation, each unit should follow the principle of flow and momentum conservation. When turbojet engine is working, it is not always maintain in steady operation. However there are different kinds of operation performance depends on various angle of throttle and all operating areas performance is showed via the Equilibrium Running Line in the Compressor Performance Map. The aforementioned all operating areas performance requires plenty of experiments to research so that it has been classified and covered by manufactures. In general, the archives, uncovering document, only show the specifications of engine designed performance, but the information of all operating areas is seldom refer in researches. Above all, the research which analyzes the specifications of engine unit operation is based on mechanical principle of turbo. By using the Microsoft Visual Basic 6.0 program design a simulator – AFSP, which can predict equilibrium running performance of turbojet engine. The sample of simulator is using J85-21B engine. First, using GasTurb 12(online test version) program complete the compressor performance map. Then using AFSP simulator predicts and proves the Equilibrium Running Line. The major result of the research is to complete the equilibrium running performance of turbojet engine prediction by using fundamental theories and design a simulator – AFSP, and therefore it accomplish the prediction of Equilibrium Running Line in limited circumstance (engine design point performance). It offers a usable method on controlling all operating areas performance of turbojet engine.
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CHOU, CHE-MING, and 周哲名. "A Neural Network-Based Simulation Study for the Performance Monitoring of a Military Turbojet Engine." Thesis, 2014. http://ndltd.ncl.edu.tw/handle/sp43uw.
Full textAbstract:
碩士國防大學理工學院
機械工程碩士班
102
In viewing to the importance of performance monitoring and trouble shooting of military engines to the readiness of air force aircrafts, the study of performance evaluation thus is of the present interests. Therefore, how to apply theorem with air-thermodynamics to an engine performance monitoring tool is valuable in researches for national defense science. To focalize, the case of an turbine engine was chosen as a typical example for the performance monitoring study, due to its popularity in modern turbine engines, while in the turbine engine performance monitoring, the artificial neural networks (ANN) algorithm was adopted in simulating the turbine engine operation mechanism, for its capability to handle highly non-linear data set with less known physics. As to the whole research works, it includes: (1) a theoretical study of the turbine engine operating characteristics, based on thermodynamics and fluid kinematics, (2) an extensive calculation study of the turbine engine performance to find the parametric analysis, (3) a comprehension study of basic ANN modeling, training and validation, (4) an ANN-based case study for the performance monitoring of a turbine engine. Through above studies, a set of two MATLAB programs was developed and used to conduct a simulation case study for the US military J85 turbojet. Major results thus obtained include the J85 turbojet engine performance was simulated by using the software of GasTurb, and assessing it’s the health state based Artificial Neural Network analysis framework. Finally, provide one kind of maintenance suggestion in academic and practice.
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Lee, Chia-Sung, and 李家崧. "Two-Dimensional Thrust-Vectoring Nozzle Design and Dynamic Model Development for a Micro Turbojet Engine and Its Stability Analysis and Verification of a Horizontal Pendulum." Thesis, 2004. http://ndltd.ncl.edu.tw/handle/72632217690701501669.
Full textAbstract:
碩士中華大學
機械與航太工程研究所
92
A two-dimensional thrust-vectoring nozzle is designed for a P-80 micro turbojet engine. The dynamic model of the side force of this thrust-vectoring engine is validated through the equilibrium of a horizontal pendulum. First, a test stand is developed to measure the axial and the lateral response of the thrust-vectoring nozzle. The performance of the nozzle is used to determine the length of the side wall. The 60mm length of the side wall gives the best steady-state response overall. The turning angle of the side wall is limited to 20 degree in order to have a reasonable exhaust temperature as well as a linear output of the side force. The maximum drop in the axial force is only 6﹪at the maximum turning angle. Then, the dynamic response of this thrust-vectoring engine is examined through a series of different frequency operation. The response of this engine by controlling the fuel flow rate can be modeled as a first order system. The time delay is 0.45 sec and the system bandwidth is 0.175Hz. The response of this engine by turning the angle of the nozzle resembles a first order system except a slight overshoot for frequency between 1.6Hz to 3.18Hz. The time delay is 0.08 sec and the system bandwidth is 2.88Hz by controlling the turning angle. The step response and the ramp response of the axial force and the side force show a good agreement between the models and the test data. The best feedback parameters are 0.05 for the proportional controller, 0.1 for the integral controller, and 1 for the derivative controller for the horizontal pendulum system. The average marginal amplitude is , the steady-state error is , and the oscillating period is 0.95 sec for the turning angle feedback. Finally, several designed parameters are investigated to reduce the marginal amplitude of the horizontal pendulum. Increasing the polar inertia reduces the effect of disturbance. Increasing the damping ratio of the bearing decreases the interference due to viscous effect but the oscillation frequency will increase. For a constant feedback turning angle of the nozzle, there is an optimal distance between the position of the side force and the pendulum support. The correction of the moment is undershoot or overshoot once deviating from the optimum length. Without the disturbance the smaller the feedback turning angle of the nozzle, the smaller the marginal amplitude is. However, the test data shows a minimal turning angle should be used to balance the horizontal pendulum. The interference from the engine itself will make the system unstable for a nozzle correction of less than turning angle. For a nozzle correction of more than turning angle, the larger the marginal amplitude is.
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Li, Jun Mo, and 李駿模. "The application of edmunds approach to full-envelope control of turbojet engines." Thesis, 1995. http://ndltd.ncl.edu.tw/handle/54173506191538997512.
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Βουγιουκλάκης, Ιωάννης. "Διερεύνηση των μηχανισμών αστοχίας επικαλύψεων δομικών στοιχείων θερμικών στροβιλομηχανών υπό συνθήκες θερμομηχανικής φόρτισης με τη χρήση μη-καταστροφικών δοκιμών." Thesis, 2003. http://nemertes.lis.upatras.gr/jspui/handle/10889/1432.
Full textAbstract:
Τo πρώτο κεφάλαιο, το οποίο είναι και το εισαγωγικό, παρουσιάζει με συντομία στον
ενδιαφερόμενο αναγνώστη τα χαρακτηριστικά, τις ιδιότητες και τις δυνατότητες των βασικών
«συστατικών» αυτής της έρευνας, που δεν είναι άλλα παρά τα υπερκράματα και οι
επικαλύψεις αυτών. Στη συνέχεια, στο επόμενο κεφάλαιο, παρουσιάζεται το βασικό
πειραματικό εργαλείο αυτής της μελέτης, δηλαδή τα πειράματα θερμομηχανικής κόπωσης
καθώς και η συγκεκριμένη πειραματική διάταξη που πρόκειται να χρησιμοποιηθεί. Το κεφάλαιο
αυτό θεωρείται ιδιαίτερα σημαντικό καθώς ασχολείται με τις βασικές παραμέτρους που
ελέγχουν τα πειράματα αυτά, καθώς και με τους μηχανισμούς αστοχίας που αναμένεται να
ενεργοποιούνται κατά τη διάρκειά τους. Το τρίτο κεφάλαιο μπορεί να θεωρηθεί και αυτό
καθαρά εισαγωγικό, όπως και το πρώτο, καθώς ασχολείται με τις βασικές αρχές και ιδιότητες
της μεθόδου της Ακουστικής Εκπομπής. Το κεφάλαιο αυτό κλείνει με μια περιγραφή του
συστήματος της Ακουστικής Εκπομπής που χρησιμοποιήθηκε στη μελέτη αυτή. Τα τρία αυτά κεφάλαια, είναι ανεξάρτητα μεταξύ τους και κρίθηκε ότι ήταν απαραίτητο να είναι όσο πιο
δυνατό πλήρη καθώς αναφέρονται σε υλικά, ιδιότητες και μεθόδους αρκετά «νέες» και μη
ευρέως γνωστές.
Η εργασία συνεχίζει στο τέταρτο κεφάλαιο με την παρουσίαση της πρώτης πειραματικής σειράς
που ασχολείται με ισόθερμα εφελκυστικά πειράματα υλικών στροβιλοκινητήρων, στη
θερμοκρασιακή περιοχή λειτουργίας τους. Με την παράλληλη εφαρμογή της μεθόδου της
Ακουστικής Εκπομπής γίνεται προσπάθεια ανίχνευσης και τελικά ανάλυσης των
ενεργοποιημένων μηχανισμών αστοχίας καθώς και εντοπισμού των βασικών παραμέτρων
απόδοσης των υλικών. Το αναλυτικό μέρος της εργασίας ολοκληρώνεται στο πέμπτο
κεφάλαιο, όπου και παρουσιάζονται δύο ανεξάρτητες σειρές πειραμάτων θερμομηχανικής
κόπωσης. Στα πειράματα αυτά, προηγείται μια προσπάθεια κατανόησης και ανίχνευσης των
μηχανισμών αστοχίας. Στη συνέχεια με την αξιοποίηση των καταγεγραμμένων σημάτων
Ακουστικής Εκπομπής επιχειρείται συσχέτιση αυτών των μηχανισμών βλάβης – αστοχίας με
ομάδες σημάτων ΑΕ καθώς και η ποσοτικοποίηση της εξέλιξης της βλάβης και του βαθμού της
υποβάθμισης του υλικού με την εξέταση της υπογραφής των σημάτων αυτών. Η επεξεργασία
και παρουσίαση των αποτελεσμάτων αυτών επιχειρήθηκε να γίνει με τέτοιο τρόπο, που να
διασφαλίζει την ανεξαρτησία των παρατηρήσεων και συμπερασμάτων για κάθε πειραματική
σειρά αλλά συγχρόνως να εξασφαλίζει και μια συνέχεια των παρατηρήσεων και των
εφαρμογών.
Η εργασία ολοκληρώνεται στο έκτο κεφάλαιο, με μια συζήτηση πάνω στην ανάλυση αυτών
των αποτελεσμάτων, όπου διατυπώνονται οι καινοτομίες της έρευνας καθώς και οι προοπτικές
εφαρμογής της σε πιο γενικευμένη και επίσημη βάση και το τί η επιστημονική κοινότητα
καθώς και τα στελέχη των σχετικών βιομηχανιών θα μπορούσαν να αναμένουν από αυτή.-
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De, Santanu. "Modeling And Computation Of Turbulent Nonreacting And Reacting Sprays." Thesis, 2011. http://etd.iisc.ernet.in/handle/2005/2379.
Full textAbstract:
Numerical modeling of several turbulent nonreacting and reacting spray jets is carried out using a fully stochastic separated flow (FSSF) approach. As is widely used, the carrier-phase is considered in an Eulerian framework, while the dispersed phase is tracked in a Lagrangian framework following the stochastic separated flow (SSF) model. Various interactions between the two phases are taken into account by means of two-way coupling. Spray evaporation is described using a thermal model with an infinite conductivity in the liquid phase. The gas-phase turbulence terms are closed using the k-� model. In the classical SSF (CSSF) approach the effects of turbulent velocity fluctuations of the gas-phase are modeled stochastically to obtain instantaneous gas-phase velocity, which subsequently is used to estimate droplet dispersion and interphase transport rates. However, in the CSSF model, no such effort is made to model the effects of the fluctuations in the gas-phase reactive scalars, namely temperature and species mass fractions. Instead, the mean value of these scalars is used while solving for the droplet governing equations and estimating various interphase source terms. Also, in flamelet model and conditional moment closure (CMC) applications of turbulent sprays, the mixture fraction is defined using conventional definition, which is no longer a conserved quantity due to associated phase change.
Therefore, in this thesis a novel mixture fraction based FSSF approach is used to stochastically model the fluctuating temperature and composition of the gas phase. These gas-phase reactive scalars are then used to refine the estimates of the heat and mass transfer rates between the droplets and the surrounding gas-phase. It is assumed that the fluctuations in the gas-phase reactive scalars are inherently associated with the fluctuation of a single conserved scalar, namely instantaneous mixture fraction. Instantaneous value of the gas-phase reactive scalars seen by individual droplets is then estimated from the instantaneous gas-phase mixture fraction, which is obtained as the Weiner process by randomly sampling a known beta-function probability density function (PDF) of the local mixture fraction field. Finally, Favre mean value of the gas-phase scalars are recovered as appropriate moments of the PDF. The present definition of the mixture fraction based on its instantaneous value facilitate exact calculation of the source terms in the transport equation for variance of the mixture fraction, whereas conventional definition leads to terms which require further modeling and simplifications. The present FSSF model also accounts for the possibility of existence of an envelope flame between the droplet and the bulk gas-phase, which greatly increases the heat and mass transfer rates to the droplet. The present model allows us to treat the occurrence of envelope flame separately which is otherwise neglected in the conventional spray combustion models.
The FSSF model is implemented into a numerical code, and different well-defined nonreacting and reacting turbulent spray jets are investigated. For the reacting spray jets, single-step irreversible reaction with infinitely fast chemistry is assumed in the body of the flow. In such cases special care must be taken with modeling the upstream boundary condition. This is because the flow from the spray jet nozzle is unreacted and yet it becomes well reacted shortly downstream. Numerical results are compared against experimental measurements as well as with predictions using the CSSF approach. Numerical results from the FSSF and CSSF model are almost identical for the nonreacting spray jets, where the fluctuations in the gas-phase scalars are relatively low. For the reacting sprays, significant differences are found between the results of the FSSF and CSSF models for the reacting spray jets, where the fluctuations in the reactive scalars are high. The FSSF model reasonably predicts many features of the jet spray flames, such as flame length, gas-phase temperature, and spray droplet velocity/diameter distribution; results appear to be close to the experimental measurements.
Finally, the combustion characteristics of the reacting spray jets are studied following classical group combustion theory. It shows that these spray jets have external group combustion mode near the nozzle-exit. Transition to internal group combustion takes place at different downstream locations based on the droplet loading and equivalence ratio at the nozzle-exit, whereas single droplet combustion regime is observed near the tip of the visible flame. Another alternate approach to study the combustion behavior of a cloud is proposed based on fraction of droplets having i) no envelope flame, ii) envelope flame, iii) extinguished envelope flame due to high slip velocity, iv) extinguished envelope flame due to droplet diameter being too small, v) both iii) and iv) above. Based on these, different group combustion behavior of the reacting spray jets are interpreted.