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Готові списки джерел за темами / Supersonic / hypersonic

Добірка наукової літератури з теми "Supersonic / hypersonic"

Автор: Grafiati

Опубліковано: 7 липня 2024

Оновлено: 7 липня 2024

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Статті в журналах з теми "Supersonic / hypersonic":

1

Yang, Lung-Jieh, and Chao-Kang Feng. "A Unified Asymptotic Theory of Supersonic, Transonic, and Hypersonic Far Fields." Axioms 11, no. 11 (November 19, 2022): 656. http://dx.doi.org/10.3390/axioms11110656.

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The problems of steady, inviscid, isentropic, irrotational supersonic plane flow passing a body with a small thickness ratio was solved by the linearized theory, which is a first approximation at and near the surface but fails at far fields from the body. Such a problem with far fields was solved by W.D. Hayes’ “pseudo-transonic” nonlinear theory in 1954. This far field small disturbance theory is reexamined in this study first by using asymptotic expansion theory. A systematic approach is adopted to obtain the nonlinear Burgers’ equation for supersonic far fields. We also use the similarity method to solve this boundary value problem (BVP) of the inviscid Burgers’ equation and obtain the nonlinear flow patterns, including the jump condition for the shock wave. Secondly, the transonic and hypersonic far field equations were obtained from the supersonic Burgers’ equation by stretching the coordinate in the y direction and considering an expansion of the freestream Mach number in terms of the transonic and hypersonic similarity parameters. The mathematical structures of the far fields of the supersonic, transonic, and hypersonic flows are unified to be the same. The similar far field flow patterns including the shock positions of a parabolic airfoil for the supersonic, transonic, and hypersonic flow regimes are exemplified and discussed.

2

Li, Yong Hong, Xin Wu Tang, and Wei Qun Zhou. "Aerodynamic and Numerical Study on the Influence of Spike Shapes at Mach 1.5." Advanced Materials Research 1046 (October 2014): 177–81. http://dx.doi.org/10.4028/www.scientific.net/amr.1046.177.

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Taking into account the issue of configuration or aerodynamic heating, most supersonic and hypersonic flight vehicles have to use the blunt-nosed body. However, in supersonic especially in hypersonic flow the strong bow shock ahead of the blunt nose introduces a rather high shock drag that affects the aerodynamic performance of the vehicles seriously. A spike mounted on a blunt body during its flight pushes the strong bow shock away from the body surface and forms recirculation flow with low pressure ahead of the body surface, and then decreases the drag. The drag reduction effects of spikes in high supersonic and hypersonic flow had been validated through experimental and numerical methods. In order to analyze the influence of the spike on aerodynamic characteristics at low supersonic (M=1.5) flow past blunt-nosed bodies, numerical studies were carried out which included the influence of the spike shape, the analysis of the fluid flow structures and the effect on the aerodynamic characteristics of a blunt body.

3

de Araujo Martos, João Felipe, Israel da Silveira Rêgo, Sergio Nicholas Pachon Laiton, Bruno Coelho Lima, Felipe Jean Costa, and Paulo Gilberto de Paula Toro. "Experimental Investigation of Brazilian 14-X B Hypersonic Scramjet Aerospace Vehicle." International Journal of Aerospace Engineering 2017 (2017): 1–10. http://dx.doi.org/10.1155/2017/5496527.

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The Brazilian hypersonic scramjet aerospace vehicle 14-X B is a technological demonstrator of a hypersonic airbreathing propulsion system based on the supersonic combustion (scramjet) to be tested in flight into the Earth’s atmosphere at an altitude of 30 km and Mach number 7. The 14-X B has been designed at the Prof. Henry T. Nagamatsu Laboratory of Aerothermodynamics and Hypersonics, Institute for Advanced Studies (IEAv), Brazil. The IEAv T3 Hypersonic Shock Tunnel is a ground-test facility able to produce high Mach number and high enthalpy flows in the test section close to those encountered during the flight of the 14-X B into the Earth’s atmosphere at hypersonic flight speeds. A 1 m long stainless steel 14-X B model was experimentally investigated at T3 Hypersonic Shock Tunnel, for freestream Mach numbers ranging from 7 to 8. Static pressure measurements along the lower surface of the 14-X B, as well as high-speed Schlieren photographs taken from the 5.5° leading edge and the 14.5° deflection compression ramp, provided experimental data. Experimental data was compared to the analytical theoretical solutions and the computational fluid dynamics (CFD) simulations, showing good qualitative agreement and in consequence demonstrating the importance of these methods in the project of the 14-X B hypersonic scramjet aerospace vehicle.

4

Milthorpe, J. F. "Simulation of supersonic and hypersonic flows." International Journal for Numerical Methods in Fluids 14, no. 3 (February 15, 1992): 267–88. http://dx.doi.org/10.1002/fld.1650140303.

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5

Xiao, Han-shan, Chao Ou, Hong-liang Ji, Zheng-chun He, Ning-yuan Liu, and Xian-xu Yuan. "Low-Cost and Aerodynamics-Aim Hypersonic Flight Experiment MF-1." MATEC Web of Conferences 316 (2020): 04006. http://dx.doi.org/10.1051/matecconf/202031604006.

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For increasing understanding of fundamental hypersonic phenomena, the flight test program, named MF-1, is to gather fundamental scientific and engineering data on the physics and technologies critical to future operational hypersonic flight with low-cost flight test platform, which is built on the retrofitted rockets. The MF-1 program is a hypersonic flight test program executed by China Aerodynamic Research and Development Center (CARDC). The MF-1 flight flew in December 2015. The flight focuses primarily on integration of instrumentation on the test vehicle, with application to boundary layer transition and shock interaction experiments. The MF-1 payload consists of a blunted 7°half angle cone, a cylinder and 33° flare configuration. The payload was boosted to Mach 5.32 utilizing a solid-rocket booster without control for the whole flight. The flight was fully successful, and measured transition under supersonic and hypersonic conditions. The heat flux data were given by the three-dimensional thermal identification method to discriminate transition zone. The preliminary analysis shows that the real-time flight data obtained by MF-1 are reliable and can be used to validate the transition predicting model and software. The results show that the existing model is able to predict the transition location of cone at a small angle-of-attack for supersonic or hypersonic flow. This paper describes the MF-1 mission and some general conclusions derived from the experiment.

6

Verhoff, A., and D. Stookesberry. "Prediction of inviscid supersonic/hypersonic aircraft flowfields." Journal of Aircraft 29, no. 4 (July 1992): 581–87. http://dx.doi.org/10.2514/3.46205.

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7

Huang, Wei, Jun-tao Chang, and Li Yan. "Mixing and combustion in supersonic/hypersonic flows." Journal of Zhejiang University-SCIENCE A 21, no. 8 (August 2020): 609–13. http://dx.doi.org/10.1631/jzus.a20mcsf1.

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8

Hu, Jiasen, and Arthur Rizzi. "Turbulent flow in supersonic and hypersonic nozzles." AIAA Journal 33, no. 9 (September 1995): 1634–40. http://dx.doi.org/10.2514/3.12861.

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9

James, Anthony. "Hot Property." Aerospace Testing International 2018, no. 3 (September 2018): 48–52. http://dx.doi.org/10.12968/s1478-2774(23)50116-2.

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With growing commercial interest in space, as well as in the development of supersonic and hypersonic aircraft, the unique capabilities of the Scirocco Plasma Wind Tunnel in Italy are increasingly in demand

10

Zhao, Lian Jin, Jia Lin, Jian Hua Wang, Jin Long Peng, De Jun Qu, and Lian Zhong Chen. "An Experimental Investigation on Transpiration Cooling for Supersonic Vehicle Nose Cone Using Porous Material." Applied Mechanics and Materials 541-542 (March 2014): 690–94. http://dx.doi.org/10.4028/www.scientific.net/amm.541-542.690.

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During hypersonic flight or cruise in the near space, the aerodynamic heating causes a very high temperature on the leading edge of hypersonic vehicles. Transpiration cooling has been recognized the most effective cooling technology. This paper presents an experimental investigation on transpiration cooling using liquid water as coolant for a nose cone model of hypersonic vehicles. The nose cone model consists of sintered porous material. The experiments were carried out in the Supersonic Jet Arc-heated Facility (SJAF) of China Academy of Aerospace Aerodynamics (CAAA) in Beijing. The cooling effect in the different regions of the model was analyzed, and the shock wave was exhibited. The pressure variations of the coolant injection system were continuously recorded. The aim of this work is to provide a relatively useful reference for the designers of coolant driving system in practical hypersonic vehicles.

Більше джерел

Дисертації з теми "Supersonic / hypersonic":

1

Higgins, Andrew J. "Investigation of detonation initiation by supersonic blunt bodies /." Thesis, Connect to this title online; UW restricted, 1996. http://hdl.handle.net/1773/10000.

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2

Denman, Paul Ashley. "Experimental study of hypersonic boundary layers and base flows." Thesis, Imperial College London, 1996. http://hdl.handle.net/10044/1/45466.

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This experimental study documents the development and separation of a hypersonic boundary layer produced naturally on the cold surface of a sharp slender cone. At the base of the conical forebody, the equilibrium turbulent boundary layer was allowed to separate over an axisymmetric rearward facing step to form a compressible base flow. The investigation was conducted in the Imperial College No.2 gun tunnel at a freestream Mach number of 9 and unit Reynolds numbers of 15 and 55 million. The compressible boundary layer study was carried out at both of the available freestream unit Reynolds numbers and the measured data include distributions of wall static pressure and heat transfer rate, together with profiles of pitot pressure through the boundary layer. Using the chordwise distribution of surface heat flux as a means of transition detection, the cone transition Reynolds number was found to be 5.4x10^. This result, together with that obtained from flat plate studies conducted in the same test facility, provided a ratio of cone to flat plate transition Reynolds number of 0.8. Boundary layer integral quantities and shape factors are derived from velocity profiles and in most cases the measured data extended close enough to the wall to detect the peak values of the integrands. The separated flow region formed at the base of the cone was documented only at the higher unit Reynolds number, a condition under which the approaching turbulent boundary layer was found to be close to equilibrium. The data include pitot pressure profiles recorded normal to the surface downstream of reattachment, together with wall static pressure and heat transfer rate distributions measured throughout the base flow region. Reattachment occurred approximately two step heights downstream of separation and a surface flow visualisation study indicated the existence of Taylor-Goertler type vortices, emanating from the reattachment line in the downstream direction. A simple shear layer expansion model is developed and shown to provide a favourable prediction of the measured pitot pressure profiles recorded downstream of the reattachment line. The success of this second order model implies that the dynamics of the corner expansion process, except in the immediate vicinity of the wall, is governed largely by inviscid pressure mechanisms and that the supersonic region of the boundary layer expansion is essentially isentropic.

3

Hunt, David Leslie. "An investigation of supersonic buffet using a Large Eddy Simulation." Thesis, Queen's University Belfast, 1995. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.318735.

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4

Husmeier, Frank. "Numerical Investigations of Transition in Hypersonic Flows over Circular Cones." Diss., The University of Arizona, 2008. http://hdl.handle.net/10150/196123.

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This thesis focuses on secondary instability mechanisms of high-speed boundary layers over cones with a circular cross section. Hypersonic transition investigations at Mach 8 are performed using Direct Numerical Simulations (DNS). At wind-tunnel conditions, these simulations allow for comparison with experimental measurements to verify fundamental stability characteristics.To better understand geometrical influences, flat-plate and cylindrical geometries are studied using after-shock conditions of the conical investigations. This allows for a direct comparison with the results of the sharp cone to evaluate the influence of the spanwise curvature and the cone opening angle. The ratio of the boundary-layer thickness to the spanwise radius is used to determine the importance of spanwise curvature effects. When advancing in the downstream direction the radius increaseslinearly while the boundary-layer thickness stays almost constant. Hence, spanwise curvature effects are strongest close to the nose and decrease in downstream direction. Their influences on the secondary instability mechanisms provide some rudimentary guidance in the design of future high-speed air vehicles.In experiments, blunting of the nose tip of the circular cone results in an increase in critical Reynolds number (c.f. Stetson et al. (1984)). However, once a certain threshold of the nose radius is exceeded, the critical Reynolds number decreases even to lower values than for the sharp cone. So far, conclusive explanations for this behavior could not be derived based on the available experimental data. Therefore, here DNS is used to study the effect of nose bluntness on secondary instability mechanisms in order to shed light on the underlying flow physics. To this end, three different nose tip radii are considered-the sharp cone, a small nose radius and a large nose radius. A small nose radius moves the transition on-set downstream, while for a large nose radius the so-called transition reversal is observed. Experimentalists hold influences of the entropy layer responsible but detailed numerical studies may lead to alternateconclusions.

5

Del, Rio Francesco. "Distortion mechanism in supersonic combustion ramjet engines." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2018.

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Анотація:

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.

6

Fuller, Eric James. "Experimental and computational investigation of helium injection into air at supersonic and hypersonic speeds." Diss., Virginia Tech, 1992. http://hdl.handle.net/10919/39977.

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Experiments were performed with two different helium injector models at different injector transverse and yaw angles in order to determine the mixing rate and core penetration of the injectant and the flow field total pressure losses. when gaseous injection occurs into a supersonic freestream. Tested in the Virginia Tech supersonic tunnel. with a freestream Mach number of 3.0 and conditions corresponding to a freestream Reynolds number of 5.0 x 107 1m. was a single. sonic. 5X underexpanded, helium jet at a downstream angle of 30° relative to the freestream. This injector was rotated from 0° to _28° to test the effects of injector yaw. The second model was an array of three supersonic, 5X underexpanded helium injectors with an exit Mach number of 1.7 and a transverse angle of 15°. This model was tested in the NASA Langley Mach 6.0, High Reynolds number tunnel, with freestream conditions corresponding to a Reynolds number of 5.4 x 10⁷ /m. The injector array as tested at yaw angles of 0° and -15°. Surface flow visualization showed that significant flow asymmetries were produced by injector yaw. Nanosecond exposure shadowgraph pictures were taken, showing the gaseous injection plume to be unsteady, and further studies demonstrated this unsteadiness was related to shock waves orthogonal to the injectant bow shock, that were generated at a frequency of 30 kHz. The primary data technique used, was a concentration probe which measured the molar concentration of helium in the flow field. Concentration data and other meanflow data was taken at several downstream axial stations and yielded contours of helium concentration, total pressure, Mach number, velocity, and mass flux, as well as the static properties. From these contour plots, the various mixing rates for each case were determined. The injectant mixing rates, expressed as the maximum concentration decay, and mixing distances were found to be unaffected by injector yaw, in the Mach 3.0 experiments, but were adversely affected by injector yaw in the Mach 6.0 experiments. One promising aspect of injector yaw was the that as the yaw angle was increased, lateral motion of the injectant plume became significant, and the turbulent mixing region area increased by approximately 34%. Comparisons of the 15° transverse angled injection into a Mach 6.0 flow to previous experiments with 15° injection into a Mach 3.0 freestream, demonstrated that there is a significant decrease in initial mixing, at Mach 6.0, resulting in a much longer mixing distance. From a parametric computational study of the Mach 6.0 experiments, the effects of adjacent injectors was found to decrease lateral spreading while increasing the vertical penetration of the injectant plume, and marginally increasing the injectant core decay rate. Matching of the computational results to the experimental results was best achieved when using the Baldwin-Lomax turbulence model without the Degani-Schiff modification.
Ph. D.

7

Lee, Jaewoo. "Efficient inverse methods for supersonic and hypersonic body design, with low wave drag analysis." Diss., Virginia Tech, 1991. http://hdl.handle.net/10919/37406.

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With the renewed interest in the supersonic and hypersonic flight vehicles, new inverse Euler methods are developed in these flow regimes where a space marching numerical technique is valid. In order to get a general understanding for the specification of target pressure distributions, a study of minimum drag body shapes was conducted over a Mach number range from 3 to 12. Numerical results show that the power law bodies result in low drag shapes, where the n=.69 (l/d = 3) or n=.70 (l/d = 5) shapes have lower drag than the previous theoretical results (n=.75 or n=.66 depending on the particular form of the theory). To validate the results, a numerical analysis was made including viscous effects and the effect of gas model. From a detailed numerical examination for the nose regions of the minimum drag bodies, aerodynamic bluntness and sharpness are newly defined. Numerous surface pressure-body geometry rules are examined to obtain an inverse procedure which is robust, yet demonstrates fast convergence. Each rule is analyzed and examined numerically within the inverse calculation routine for supersonic (M_∞= 3) and hypersonic (M_∞ = 6.28) speeds. Based on this analysis, an inverse method for fully three dimensional supersonic and hypersonic bodies is developed using the Euler equations. The method is designed to be easily incorporated into existing analysis codes, and provides the aerodynamic designer with a powerful tool for design of aerodynamic shapes of arbitrary cross section. These shapes can correspond to either "wing like" pressure distributions or to "body like" pressure distributions. Examples are presented illustrating the method for a non-axisymmetric fuselage type pressure distribution and a cambered wing type application. The method performs equally well for both nonlifting and lifting cases. For the three dimensional inverse procedure, the inverse solution existence and uniqueness problem are discussed. Sample calculations demonstrating this problem are also presented.
Ph. D.

8

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

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

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

9

Schreyer, Anne-Marie [Verfasser]. "Experimental investigations of supersonic and hypersonic shock wave/turbulent boundary layer interactions / Anne-Marie Schreyer." München : Verlag Dr. Hut, 2013. http://d-nb.info/1045126853/34.

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10

Rock, Christopher. "Experimental Studies of Injector Array Configurations for Circular Scramjet Combustors." Diss., Virginia Tech, 2010. http://hdl.handle.net/10919/77208.

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A flush-wall injector model and a strut injector model representative of state of the art scramjet engine combustion chambers were experimentally studied in a cold-flow (non-combusting) environment to determine their fuel-air mixing behavior under different operating conditions. The experiments were run at nominal freestream Mach numbers of 2 and 4, which simulates combustor conditions for nominal flight Mach numbers of 5 and 10. The flush-wall injector model consists of sixteen inclined, round, sonic injectors distributed around the wall of a circular duct. The strut injector model has sixteen inclined, round, sonic injectors distributed across four struts within a circular duct. The struts are slender, inclined at a low angle to minimize drag, and have two injectors on each side. The experiments investigated the effects of injectant molecular weight, freestream Mach number, and jet-to-freestream momentum flux ratio on the fuel-air mixing process. Helium, methane, and air injectants were studied to vary the injectant molecular weight over the range of 4-29. All of these experiments were performed to support the needs of an integrated experimental and computational research program, which has the goal of upgrading the turbulence models that are used for Computational Fluid Dynamics predictions of the flow inside a scramjet combustor. The primary goals of this study were to use injector models that represent state of the art scramjet engine combustion chambers to provide validation data to support the development of turbulence model upgrades and to add to the sparse database of mixing results in such configurations. The main experimental results showed that higher molecular weight injectants had approximately the same amount of penetration in the far field as lower molecular weight injectants at the same jet-to-freestream momentum flux ratio. Higher molecular weight injectants also demonstrated a mixing rate that was the same as or slower than lower molecular weight injectants depending on the flow conditions. A comparison of the experimental results for the two different injector models revealed that the flush-wall injector mixed significantly faster than the strut injector in all of the experimental cases.
Ph. D.

Більше джерел

Книги з теми "Supersonic / hypersonic":

1

Center, Langley Research, ed. Wave-interactions in supersonic and hypersonic flows. Norfolk, Va: Old Dominion University Research Foundation, 1990.

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2

Lakin, William D. Wave-interactions in supersonic and hypersonic flows. Norfolk, Va: Old Dominion University Research Foundation, 1990.

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3

Anderson, Griffin Y. An outlook on hypersonic flight. New York: AIAA, 1987.

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4

North Atlantic Treaty Organization. Advisory Group for Aerospace Research and Development. Hypersonic combined cycle propulsion. Neuilly sur Seine, France: AGARD, 1990.

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5

North Atlantic Treaty Organization. Advisory Group for Aerospace Research and Development. Hypersonic combined cycle propulsion. Neuilly sur Seine, France: AGARD, 1990.

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6

Miles, Richard B. Filtered Rayleigh scattering measurements in supersonic/hypersonic facilities. Washington: AIAA, 1992.

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7

Center, Langley Research, and United States. National Aeronautics and Space Administration. Scientific and Technical Information Division., eds. Nonparallel instability of supersonic and hypersonic boundary layers. Washington, D.C: National Aeronautics and Space Administration, Office of Management, Scientific and Technical Information Division, 1991.

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8

D, Carboni Jeanne, Supersonic investigation of two dimensional ...., and United States. National Aeronautics and Space Administration., eds. Supersonic investigation of two-dimensional hypersonic exhaust nozzles. [Washington, DC: National Aeronautics and Space Administration, 1992.

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9

D, Carboni Jeanne, Supersonic investigation of two dimensional..., and United States. National Aeronautics and Space Administration., eds. Supersonic investigation of two-dimensional hypersonic exhaust nozzles. [Washington, DC: National Aeronautics and Space Administration, 1992.

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10

L, Pittman Jimmy, and United States. National Aeronautics and Space Administration. Scientific and Technical Information Branch., eds. Aerodynamic characteristics of a distinct wing-body configuration at Mach 6: Experiment, theory, and the hypersonic isolation principle. [Washington, DC]: National Aeronautics and Space Administration, Scientific and Technical Information Branch, 1985.

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Частини книг з теми "Supersonic / hypersonic":

1

Ingenito, Antonella. "Design of Supersonic/Hypersonic Vehicles." In Subsonic Combustion Ramjet Design, 9–17. Cham: Springer International Publishing, 2021. http://dx.doi.org/10.1007/978-3-030-66881-5_3.

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2

Stemmer, Christian, and Nikolaus A. Adams. "Supersonic and Hypersonic Boundary-Layer Flows." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 77–91. Berlin, Heidelberg: Springer Berlin Heidelberg, 2009. http://dx.doi.org/10.1007/978-3-642-00262-5_4.

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3

Vervisch, P., and A. Bourdon. "The Recombination of Ionized Species in Supersonic Flows." In Molecular Physics and Hypersonic Flows, 525–42. Dordrecht: Springer Netherlands, 1996. http://dx.doi.org/10.1007/978-94-009-0267-1_35.

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4

Marcos, T. V. C., D. Romanelli Pinto, G. S. Moura, A. C. Oliveira, J. B. Chanes, P. G. P. Toro, and M. A. S. Minucci. "Supersonic Combustion Flow Visualization at Hypersonic Flow." In 28th International Symposium on Shock Waves, 1041–47. Berlin, Heidelberg: Springer Berlin Heidelberg, 2012. http://dx.doi.org/10.1007/978-3-642-25685-1_158.

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5

Wegener, Peter P. "Toward High Speed: Supersonic and Hypersonic Flight." In What Makes Airplanes Fly?, 145–66. New York, NY: Springer US, 1991. http://dx.doi.org/10.1007/978-1-4684-0403-6_10.

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6

Wegener, Peter P. "Toward High Speed: Supersonic and Hypersonic Flight." In What Makes Airplanes Fly?, 169–93. New York, NY: Springer New York, 1997. http://dx.doi.org/10.1007/978-1-4612-2254-5_10.

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7

Smits, Alexander J., and M. Pino Martin. "Turbulence in Supersonic and Hypersonic Boundary Layers." In Solid mechanics and its applications, 221–30. Dordrecht: Springer Netherlands, 2006. http://dx.doi.org/10.1007/978-1-4020-4150-1_21.

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8

Verhoff, A., D. C. Stookesberry, B. M. Hopping, and T. R. Michal. "Supersonic/Hypersonic Euler Flowfield Prediction Method for Aircraft Configurations." In Numerical and Physical Aspects of Aerodynamic Flows IV, 189–204. Berlin, Heidelberg: Springer Berlin Heidelberg, 1990. http://dx.doi.org/10.1007/978-3-662-02643-4_12.

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9

Romanelli Pinto, D., T. V. C. Marcos, R. L. M. Alcaide, A. C. Oliveira, J. B. Chanes, P. G. P. Toro, and M. A. S. Minucci. "Supersonic Combustion Experimental Investigation at T2 Hypersonic Shock Tunnel." In 28th International Symposium on Shock Waves, 1049–55. Berlin, Heidelberg: Springer Berlin Heidelberg, 2012. http://dx.doi.org/10.1007/978-3-642-25685-1_159.

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10

Seror, S., and L. Kosarev. "Turbulence Compressibility Effects for Supersonic and Hypersonic Separated Flows." In 30th International Symposium on Shock Waves 1, 263–67. Cham: Springer International Publishing, 2017. http://dx.doi.org/10.1007/978-3-319-46213-4_43.

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Тези доповідей конференцій з теми "Supersonic / hypersonic":

1

MASON, W., and JAEWOO LEE. "On optimal supersonic/hypersonic bodies." In Flight Simulation Technologies Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1990. http://dx.doi.org/10.2514/6.1990-3072.

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2

SOBIECZKY, H. "Generic supersonic and hypersonic configurations." In 9th Applied Aerodynamics Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1991. http://dx.doi.org/10.2514/6.1991-3301.

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3

YOON, W., and T. CHUNG. "Numerical studies of supersonic/hypersonic combustion." In 30th Aerospace Sciences Meeting and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1992. http://dx.doi.org/10.2514/6.1992-94.

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4

Landon, Mark, Darryl Hall, Jerry Udy, and Ernest Perry. "Automatic supersonic/hypersonic aerodynamic shape optimization." In 12th Applied Aerodynamics Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1994. http://dx.doi.org/10.2514/6.1994-1898.

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5

Abouali, Omid, and Goodarz Ahmadi. "Bow Shock Effect on Particle Transport and Deposition in a Hypersonic Impactor." In ASME/JSME 2003 4th Joint Fluids Summer Engineering Conference. ASMEDC, 2003. http://dx.doi.org/10.1115/fedsm2003-45072.

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Анотація:

Supersonic/hypersonic impactors are used as a collector and/or size separator of nano- and micro-particles. Thin film and fine line pattern deposition by aerosol jets are other applications of deposition of supersonic/hypersonic impactors. At extremely low backpressures, the exiting flow from a nozzle forms a supersonic free jet. The supersonic jet forms a strong normal shock in the front of the impactor plate. The stagnation pressure, backpressure and distance between the nozzle exit and the impactor plate affect the flow field. Due to the rather complicated flow in the impactor, studies of particle motions in supersonic impactors are rather scarce. In this study the airflow and particle transport and deposition in a supersonic/hypersonic impactor are numerically simulated. The axisymmetric compressible Navier-Stokes equation is solved and the flow properties are evaluated. It is assumed that the particle concentration is dilute, to the extent that the presence of particles does not alter the flow field. Deposition of different size particles under different operating conditions is studied. The importance of drag, lift and Brownian forces on particle motions in supersonic impactors is discussed. Sensitivity of the simulation results to the use of different expressions for the drag force is also examined. A strong bow shock on the flowfield has much effect in drag forces on particles. It is shown that the Stokes-Cunningham drag with variable correction coefficient is most suitable for computer simulation studies of nano-particles in supersonic/hypersonic impactors. The computer simulation results are shown to compare favorably with the experimental data.

6

Zhang, Chen-an, Zheng-yin Ye, and Wei-wei Zhang. "Aeroservoelastic Analysis for Supersonic and Hypersonic Missiles." In 45th AIAA Aerospace Sciences Meeting and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2007. http://dx.doi.org/10.2514/6.2007-1073.

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7

NICKERSON, G., S. DUNN, and D. MIGDAL. "Optimized supersonic exhaust nozzles for hypersonic propulsion." In 24th Joint Propulsion Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1988. http://dx.doi.org/10.2514/6.1988-3161.

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8

Arai, Takakage, Sosuke Sugano, and Shoji Sakaue. "Interaction between Supersonic Cavity Flow and Streamwise Vortices for Supersonic Mixing Enhancement." In 20th AIAA International Space Planes and Hypersonic Systems and Technologies Conference. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2015. http://dx.doi.org/10.2514/6.2015-3613.

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9

Thakur, Amit, and Corin Segal. "Flameholding Analyses in Supersonic Flow." In 12th AIAA International Space Planes and Hypersonic Systems and Technologies. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2003. http://dx.doi.org/10.2514/6.2003-6909.

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10

Abouali, Omid, and Goodarz Ahmadi. "Numerical Modeling of Upstream Nozzle Effect in Supersonic/Hypersonic Impactors for Nano-Particles." In ASME 2005 Fluids Engineering Division Summer Meeting. ASMEDC, 2005. http://dx.doi.org/10.1115/fedsm2005-77433.

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Анотація:

In this study the performance of supersonic and hypersonic impactors under various operating conditions was analyzed using a computer simulation model. The study was focused on the effect of the nozzle upstream condition on the performance of the supersonic and hypersonic impactors. In our earlier work, the computational domain covered downstream of the nozzle with a sonic boundary condition at the inlet. In the present study, the computational domain included the upstream nozzle where the flow and particles enter with at low velocities. Axisymmetric forms of the compressible Navier-Stokes and energy equations were solved and the gas flow and thermal condition in the impactor were for evaluated. A Lagrangian particle trajectory analysis procedure was used and the deposition rates of different size particles under various operating conditions were studied. For dilute particle concentrations, one-way interaction was assumed and the effect of particles on gas flow field was ignored. The importance of drag and Brownian forces on particle motions in supersonic/hypersonic impactors was analyzed. Sensitivity of the simulation results to the use of different expressions for the drag force was also examined. It was shown that when the upstream nozzle is included in the computational model, the Stokes-Cunningham drag with variable correction coefficient and a constant Cunningham correction factor based on stagnation point properties lead to the same results. Thus these drag laws are most suitable for computer simulation studies of nano-particles in supersonic/hypersonic impactors. The computer simulation results were shown to compare favorably with the experimental data.

Звіти організацій з теми "Supersonic / hypersonic":

1

Kostoff, Ronald N., Henry J. Eberhart, and Darrell R. Toothman. Science and Technology Text Mining: Hypersonic and Supersonic Flow. Fort Belvoir, VA: Defense Technical Information Center, November 2003. http://dx.doi.org/10.21236/ada418717.

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2

Herbert, Thorwald. Stability of Boundary Layers at High Supersonic and Hypersonic Speeds. Fort Belvoir, VA: Defense Technical Information Center, May 1992. http://dx.doi.org/10.21236/ada250900.

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3

Lempert, Walter R., and Richard B. Miles. Quantitative Imaging of Time-Evolving Structure for Supersonic and Hypersonic Flows. Fort Belvoir, VA: Defense Technical Information Center, March 1995. http://dx.doi.org/10.21236/ada297721.

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