Dissertations / Theses on the topic 'Supersonic / hypersonic'
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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.
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.
Abstract:
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.
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.
Abstract:
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.
Abstract:
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.
Abstract:
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.
Abstract:
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.
Abstract:
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.
10
Rock, Christopher. "Experimental Studies of Injector Array Configurations for Circular Scramjet Combustors." Diss., Virginia Tech, 2010. http://hdl.handle.net/10919/77208.
Abstract:
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.
11
Smith, Theodore Brooke. "Development and Ground Testing of Direct Measuring Skin Friction Gages for High Enthalpy Supersonic Flight Tests." Diss., Virginia Tech, 2001. http://hdl.handle.net/10919/29351.
Abstract:
A series of direct-measuring skin friction gages were developed for a high-speed, high-temperature environment of the turbulent boundary layer in flows such as that in supersonic combustion ramjet (scramjet) engines, with a progression from free-jet ground tests to a design for an actual hypersonic scramjet-integrated flight vehicle. The designs were non-nulling, with a sensing head that was flush with the model wall and surrounded by a small gap. Thus, the shear force due to the flow along the wall deflects the head, inducing a measurable strain. Strain gages were used to detect the strain. The gages were statically calibrated using a direct force method. The designs were verified by testing in a well-documented Mach 2.4 cold flow. Results of the cold-flow tests were repeatable and within 15% of the value of Cf estimated from simple theory. The first gage design incorporated a cantilever beam with semiconductor strain gages to sense the shear on the floating head. Cooling water was routed both internally and around the external housing in order to control the temperature of the strain gages. This first gage was installed and tested in a rocket-based-combined-cycle (RBCC) engine model operating in the scramjet mode. The free-jet facility provided a Mach 6.4 flow with P0 = 1350 psia (9310 kPa) and T0 = 2800 °R (1555 °K). Local wall temperatures were measured between 850 and 900 °R (472-500 °K). Output from the RBCC scramjet tests was reasonable and repeatable. A second skin friction gage was designed for and tested in a wind tunnel model of the Hyper-X flight vehicle scramjet engine. These unsuccessful tests revealed the need for a radically different skin friction gage design. The third and final skin friction gage was specifically developed to be installed on the Hyper-X flight vehicle. Rather than the cantilever beam and semiconductor strain gages, the third skin friction gage made use of a flexure ring and metal foil strain gages to sense the shear. The water-cooling and oil-fill used on the previous skin friction sensors were eliminated. It was qualified for flight through a rigorous series of environmental tests, including pressure, temperature, vibration, and heat flux tests. Finally, the third skin friction gage was tested in the Hyper-X Engine Model (HXEM), a full-scale-partial-width wind tunnel model of the flight vehicle engine. These tests were conducted at Mach 6.5 enthalpy with P0 = 555 psia (3827 kPa) and h0 = 900 Btu/lbm in a freejet facility. The successful testing in the wind tunnel scramjet model provided the final verification of the gage before installation in the flight vehicle engine. The development, testing, and results of all three skin friction gages are discussed.Ph. D.
12
Witt, Michael A. "Investigation in the feasibility of using solid fuel ramjets for high supersonic/low hypersonic tactical missiles." Thesis, Monterey, California. Naval Postgraduate School, 1989. http://hdl.handle.net/10945/26015.
13
Axdahl, Erik Lee. "A study of premixed, shock-induced combustion with application to hypervelocity flight." Diss., Georgia Institute of Technology, 2013. http://hdl.handle.net/1853/50290.
Abstract:
One of the current goals of research in hypersonic, airbreathing propulsion is access to higher Mach numbers. A strong driver of this goal is the desire to integrate a scramjet engine into a transatmospheric vehicle airframe in order to improve performance to low Earth orbit (LEO) or the performance of a semi-global transport. An engine concept designed to access hypervelocity speeds in excess of Mach 10 is the shock-induced combustion ramjet (i.e. shcramjet). This dissertation presents numerical studies simulating the physics of a shcramjet vehicle traveling at hypervelocity speeds with the goal of understanding the physics of fuel injection, wall autoignition mitigation, and combustion instability in this flow regime.
This research presents several unique contributions to the literature. First, different classes of injection are compared at the same flow conditions to evaluate their suitability for forebody injection. A novel comparison methodology is presented that allows for a technically defensible means of identifying outperforming concepts. Second, potential wall cooling schemes are identified and simulated in a parametric manner in order to identify promising autoignition mitigation methods. Finally, the presence of instabilities in the shock-induced combustion zone of the flowpath are assessed and the analysis of fundamental physics of blunt-body premixed, shock-induced combustion is accelerated through the reformulation of the Navier Stokes equations into a rapid analysis framework. The usefulness of such a framework for conducting parametric studies is demonstrated.
14
Malo-Molina, Faure Joel. "Numerical study of innovative scramjet inlets coupled to combustors using hydrocarbon-air mixture." Diss., Georgia Institute of Technology, 2010. http://hdl.handle.net/1853/33906.
Abstract:
To advance the design of hypersonic vehicles, high-fidelity multi-physics CFD is used to characterize 3-D scramjet flow-fields in two novel streamline traced configurations. The two inlets, Jaws and Scoop, are analyzed and compared to a traditional rectangular inlet used as a baseline for on/off-design conditions. The flight trajectory conditions selected are Mach 6 and a dynamic pressure of 1,500 psf (71.82 kPa). Analysis of these hypersonic inlets is performed to investigate distortion effects downstream with multiple single cavity combustors acting as flame holders, and several fuel injection strategies. The best integrated scramjet inlet/combustor design is identified. The flow physics is investigated and the integrated performance impact of the two innovative scramjet inlet designs is quantified. Frozen and finite rate chemistry is simulated with 13 gaseous species and 20 reactions for an Ethylene/air finite-rate chemical model. In addition, URANS and LES modeling are compared to explore overall flow structure and to contrast individual numerical methods.
The flow distortion in Jaws and Scoop is similar to some of the distortion in the traditional rectangular inlet, despite design differences. The baseline and Jaws performance attributes are stronger than Scoop, but Jaws accomplishes this while eradicating the cowl lip interaction, and lessening the total drag and spillage penalties.
The innovative inlets work best on-design, whereas for off-design, the traditional inlet is best. Early pressure losses and flow distortions in the isolator aid the mixing of air and fuel, and improve the overall efficiency of the system. Although the trends observed with and without chemical reactions are similar, the former yields roughly 10% higher mixing efficiency and upstream reactions are present. These show a significant impact on downstream development. Unsteadiness in the combustor increases the mixing efficiency, varying the flame anchoring and combustion pressure effects upstream of the step.
15
Laible, Andreas Christian. "Numerical Investigation of Boundary-Layer Transition for Cones at Mach 3.5 and 6.0." Diss., The University of Arizona, 2011. http://hdl.handle.net/10150/205419.
Abstract:
Transition in high-speed boundary layers is investigated using direct numerical simulation (DNS). A compressible Navier-Stokes code that is specifically tailored towards accurate and efficient simulations of boundary layer stability and boundary layer transition was developed and thoroughly validated. Particular emphasis was put into the adoption of a high-order accurate spatial discretization including a boundary closure with the same stencil width as the interior scheme. Oblique breakdown has been shown, using both temporal and spatial DNS, to be a viable route to transition for the boundary layer of the sharp 7° cone at Mach 3.5 investigated by Corke 2002. A 'wedge-shaped' transitional regime was observed to be characteristic for this type of breakdown on the cone geometry. Furthermore, it was shown that the dominance of the longitudinal mode in the nonlinear transition regime of oblique breakdown is due to a continuously nonlinear forced transient growth. That is the primary pair of oblique waves permanently 'seeds' disturbances into the longitudinal mode, where these disturbances exhibit non-modal unstable behavior. In addition to the simulations of controlled transition via oblique breakdown, six simulations have been conducted and analyzed where transition is initiated by multiple primary waves. Despite the broader spectrum of primary waves, typical features of oblique breakdown are still apparent in these simulations and therefore, it may be conjectured, that oblique breakdown initiated by one primary pair of waves is a good model for the nonlinear processes in natural transition. Furthermore, hypersonic boundary layer stability and transition for a flared and a straight cone at Mach 6 was investigated. In particular, a comparative investigation between both geometries regarding the K-type breakdown was performed in order to give some indications towards the open question how strong the nonlinear transition processis altered by the cone flare.
16
Retaureau, Ghislain J. "On recessed cavity flame-holders in supersonic cross-flows." Diss., Georgia Institute of Technology, 2012. http://hdl.handle.net/1853/43703.
Abstract:
Flame-holding in a recessed cavity is investigated experimentally in a Mach 2.5 preheated cross-flow for both stable and unstable combustion, with a relatively low preheating. Self-sustained combustion is investigated for stagnation pressures and temperatures reaching 1.4 MPa and 750 K. In particular, cavity blowout is characterized with respect to cavity aspect ratio (L/D =2.84 - 3.84), injection strategy (floor - ramp), aft ramp angle (90 deg - 22.5 deg) and multi-fuel mixture (CH₄-H₂ or CH₄-C₂H₄ blends). The results show that small hydrogen addition to methane leads to significant increase in flame stability, whereas ethylene addition has a more gradual effect. Since the multi-fuels used here are composed of a slow and a fast chemistry fuel, the resulting blowout region has a slow (methane dominant) and a fast (hydrogen or ethylene dominant) branch. Regardless of the fuel composition, the pressure at blowout is close to the non-reacting pressure imposed by the cross-flow, suggesting that combustion becomes potentially unsustainable in the cavity at the sub-atmospheric pressures encountered in these supersonic studies. The effect of preheating is also investigated and results show that the stability domain broadens with increasing stagnation temperature. However, smaller cavities appear less sensitive to the cross-flow preheating, and stable combustion is achieved over a smaller range of fuel flow rate, which may be the result of limited residence and mixing time. The blowout data point obtained at lower fuel flow rate fairly matches the empirical model developed by Rasmussen et al. for floor injection phi = 0.0028 Da^-.8, where phi is the equivalence ratio and Da the Damkohler number. An alternate model is proposed here that takes into account the ignition to scale the blowout data. Since the mass of air entrained into the cavity cannot be accurately estimated and the cavity temperature is only approximated from the wall temperature, the proposed scaling has some uncertainty. Nevertheless the new phi-Da scaling is shown to preserve the subtleties of the blowout trends as seen in the current experimental data.
17
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.
18
Griffiths, Alan David, and alan griffiths@anu edu au. "Development and demonstration of a diode laser sensor for a scramjet combustor." The Australian National University. Faculty of Science, 2005. http://thesis.anu.edu.au./public/adt-ANU20051114.132736.
Abstract:
Hypersonic vehicles, based on scramjet engines, have the potential to deliver inexpensive access to space when compared with rocket propulsion. The technology, however, is in its infancy and there is still much to be learned from fundamental studies.¶
Flows that represent the conditions inside a scramjet engine can be generated in ground tests using a free-piston shock tunnel and a combustor model. These facilities provide a convenient location for fundamental studies and principles learned during ground tests can be applied to the design of a full-scale vehicle.¶
A wide range of diagnostics have been used for studying scramjet flows, including surface measurements and optical visualisation techniques.¶
The aim of this work is to test the effectiveness of tunable diode laser absorption spectroscopy (TDLAS) as a scramjet diagnostic.¶
TDLAS utilises the spectrally narrow emission from a diode laser to probe individual
absorption lines of a target species. By varying the diode laser injection
current, the laser emission wavelength can be scanned to rapidly obtain a profile of the spectral line. TDLAS has been used previously for gas-dynamic sensing applications and, in the configuration used in this work, is sensitive to
temperature and water vapour concentration.¶
The design of the sensor was guided by previous work. It incorporated aspects of designs that were considered to be well suited to the present application. Aspects of the design which were guided by the literature included the laser emission wavelength, the use of fibre optics and the detector used. The laser emission wavelength was near 1390 nm to coincide with relatively strong water vapour transitions. This wavelength allowed the use of telecommunications optical fibre and components for light delivery. Detection used a dual-beam, noise cancelling detector.¶
The sensor was validated before deployment in a low-pressure test cell and a hydrogenair flame. Temperature and water concentration measurements were verified to within 5% up to 1550 K. Verification accuracy was limited by non-uniformity along the beam path during flame measurements.¶
Measurements were made in a scramjet combustor operating in a flow generated by the T3 shock tunnel at the Australian National University. Within the scramjet combustor, hydrogen was injected into a flame-holding cavity and the sensor was operated downstream in the expanded, supersonic, post-combustion flow. The sensor was operated at a maximum repetition rate of 20 kHz and
could resolve variation in temperature and water concentration over the 3ms running time of the facility.¶
Results were repeatable and the measurement uncertainty was smaller than the turbulent fluctuations in the flow. The scramjet was operated at two fuel-lean equivalence ratios and the sensor was able to show differences between the two operating conditions. In addition, vertical traversal of the sensor revealed variation in flow conditions across the scramjet duct.¶
The effectiveness of the diagnostic was tested by comparing results with those from other measurement techniques, in particular pressure and OH fluorescence measurements, as well as comparison with computational simulation.¶
Combustion was noted at both of the tested operating conditions in data from all three measurement techniques.¶
Computation simulation of the scramjet flow significantly under-predicted the water vapour concentration. The discrepancy between experiments and simulation was not apparent in either the pressure measurements or the OH fluorescence, but was clear in the diode laser results.¶
The diode laser sensor, therefore, was able to produce quantitative results which were useful for comparison with a CFD model of the scramjet and were complimentary to information provided by other diagnostics.
19
Bonanos, Aristides Michael. "Scramjet Operability Range Studies of an Integrated Aerodynamic-Ramp-Injector/Plasma-Torch Igniter with Hydrogen and Hydrocarbon Fuels." Diss., Virginia Tech, 2005. http://hdl.handle.net/10919/28847.
Abstract:
An integrated aerodynamic-ramp-injector/plasma-torch-igniter of original design was tested in a Mâ = 2, unvitiated, heated flow facility arranged as a diverging duct scramjet combustor. The facility operated at a total temperature of 1000 K and total pressure of 330 kPa. Hydrogen (H2), ethylene (C2H4) and methane (CH4) were used as fuels, and a wide range of global equivalence ratios were tested. The main data obtained were wall static pressure measurements, and the presence of combustion was determined based on the pressure rises obtained.
Supersonic and dual-mode combustion were achieved with hydrogen and ethylene fuel, whereas very limited heat release was obtained with the methane. Global operability limits were determined to be 0.07 < Ï < 0.31 for hydrogen, and 0.14 < Ï < 0.48 for ethylene. The hydrogen fuel data for the aeroramp/torch system was compared to data from a physical 10º unswept compression ramp injector and similar performance was found with the two arrangements. With hydrogen and ethylene as fuels and the aeroramp/plasma-torch system, the effect of varying the air total temperature was investigated. Supersonic combustion was achieved with temperatures as low as 530K and 680K for the two fuels, respectively. These temperatures are facility/operational limits, not combustion limits.
The pressure profiles were analyzed using the Ramjet Propulsion Analysis (RJPA) code. Results indicate that both supersonic and dual-mode ramjet combustion were achieved. Combustion efficiencies varied with Ï from a high of about 75% to a low of about 45% at the highest Ï . With a theoretical diffuser and nozzle assumed for the configuration and engine, thrust was computed for each fuel. Fuel specific impulse was on average 3000 and 1000 seconds for hydrogen and ethylene respectively, and air specific impulse varied from a low of about 9 sec to a high of about 24 sec (for both fuels) for the To = 1000K test condition.
The GASP RANS code was used to numerically simulate the injection and mixing process of the fuels. The results of this study were very useful in determining the suitability of the selected plasma torch locations. Further, this tool can be used to determine whether combustion is theoretically possible or not.Ph. D.
20
Morham, Brett G. "Numerical Examination of Flow Field Characteristics and Fabri Choking of 2D Supersonic Ejectors." DigitalCommons@CalPoly, 2010. https://digitalcommons.calpoly.edu/theses/340.
Abstract:
An automated computer simulation of the two-dimensional planar Cal Poly Supersonic Ejector test rig is developed. The purpose of the simulation is to identify the operating conditions which produce the saturated, Fabri choke and Fabri block aerodynamic flow patterns. The effect of primary to secondary stagnation pressure ratio on the efficiency of the ejector operation is measured using the entrainment ratio which is the secondary to primary mass flow ratio.
The primary flow of the ejector is supersonic and the secondary (entrained) stream enters the ejector at various velocities at or below Mach 1. The primary and secondary streams are both composed of air. The primary plume boundary and properties are solved using the Method of Characteristics. The properties within the secondary stream are found using isentropic relations along with stagnation conditions and the shape of the primary plume. The solutions of the primary and secondary streams iterate on a pressure distribution of the secondary stream until a converged solution is attained. Viscous forces and thermo-chemical reactions are not considered.
For the given geometry the saturated flow pattern is found to occur below stagnation pressure ratios of 74. The secondary flow of the ejector becomes blocked by the primary plume above pressure ratios of 230. The Fabri choke case exists between pressure ratios of 74 and 230, achieving optimal operation at the transition from saturated to Fabri choked flow, near the pressure ratio of 74. The case of optimal expansion yields an entrainment ratio of 0.17. The entrainment ratio results of the Cal Poly Supersonic Ejector simulation have an average error of 3.67% relative to experimental data. The accuracy of this inviscid simulation suggests ejector operation in this regime is governed by pressure gradient rather than viscous effects.
21
Benyo, Theresa Louise. "Analytical and computational investigations of a magnetohydrodynamics (MHD) energy-bypass system for supersonic gas turbine engines to enable hypersonic flight." Thesis, Kent State University, 2014. http://pqdtopen.proquest.com/#viewpdf?dispub=3618922.
Abstract:
Historically, the National Aeronautics and Space Administration (NASA) has used rocket-powered vehicles as launch vehicles for access to space. A familiar example is the Space Shuttle launch system. These vehicles carry both fuel and oxidizer onboard. If an external oxidizer (such as the Earth's atmosphere) is utilized, the need to carry an onboard oxidizer is eliminated, and future launch vehicles could carry a larger payload into orbit at a fraction of the total fuel expenditure. For this reason, NASA is currently researching the use of air-breathing engines to power the first stage of two-stage-to-orbit hypersonic launch systems. Removing the need to carry an onboard oxidizer leads also to reductions in total vehicle weight at liftoff. This in turn reduces the total mass of propellant required, and thus decreases the cost of carrying a specific payload into orbit or beyond. However, achieving hypersonic flight with air-breathing jet engines has several technical challenges. These challenges, such as the mode transition from supersonic to hypersonic engine operation, are under study in NASA's Fundamental Aeronautics Program.
One propulsion concept that is being explored is a magnetohydrodynamic (MHD) energy- bypass generator coupled with an off-the-shelf turbojet/turbofan. It is anticipated that this engine will be capable of operation from takeoff to Mach 7 in a single flowpath without mode transition. The MHD energy bypass consists of an MHD generator placed directly upstream of the engine, and converts a portion of the enthalpy of the inlet flow through the engine into electrical current. This reduction in flow enthalpy corresponds to a reduced Mach number at the turbojet inlet so that the engine stays within its design constraints. Furthermore, the generated electrical current may then be used to power aircraft systems or an MHD accelerator positioned downstream of the turbojet. The MHD accelerator operates in reverse of the MHD generator, re-accelerating the exhaust flow from the engine by converting electrical current back into flow enthalpy to increase thrust. Though there has been considerable research into the use of MHD generators to produce electricity for industrial power plants, interest in the technology for flight-weight aerospace applications has developed only recently.
In this research, electromagnetic fields coupled with weakly ionzed gases to slow hypersonic airflow were investigated within the confines of an MHD energy-bypass system with the goal of showing that it is possible for an air-breathing engine to transition from takeoff to Mach 7 without carrying a rocket propulsion system along with it. The MHD energy-bypass system was modeled for use on a supersonic turbojet engine. The model included all components envisioned for an MHD energy-bypass system; two preionizers, an MHD generator, and an MHD accelerator. A thermodynamic cycle analysis of the hypothesized MHD energy-bypass system on an existing supersonic turbojet engine was completed. In addition, a detailed thermodynamic, plasmadynamic, and electromagnetic analysis was combined to offer a single, comprehensive model to describe more fully the proper plasma flows and magnetic fields required for successful operation of the MHD energy bypass system.
The unique contribution of this research involved modeling the current density, temperature, velocity, pressure, electric field, Hall parameter, and electrical power throughout an annular MHD generator and an annular MHD accelerator taking into account an external magnetic field within a moving flow field, collisions of electrons with neutral particles in an ionized flow field, and collisions of ions with neutral particles in an ionized flow field (ion slip). In previous research, the ion slip term has not been considered.
The MHD energy-bypass system model showed that it is possible to expand the operating range of a supersonic jet engine from a maximum of Mach 3.5 to a maximum of Mach 7. The inclusion of ion slip within the analysis further showed that it is possible to 'drive' this system with maximum magnetic fields of 3 T and with maximum conductivity levels of 11 mhos/m. These operating parameters better the previous findings of 5 T and 10 mhos/m, and reveal that taking into account collisions between ions and neutral particles within a weakly ionized flow provides a more realistic model with added benefits of lower magnetic fields and conductivity levels especially at the higher Mach numbers. (Abstract shortened by UMI.)
22
Moreira, Farney Coutinho. "On the behavior of upwind schemes applied to three-dimensional supersonic and hypersonic cold gas flow simulations of aerospace configurations." Instituto Tecnológico de Aeronáutica, 2007. http://www.bd.bibl.ita.br/tde_busca/arquivo.php?codArquivo=397.
Abstract:
The present work describes the efforts towards the implementation of upwind schemes to simulate supersonic and hypersonic cold gás flows. The class of flux vector splitting schemes has been chosen, and the particular methods implemented are the van Leer and Liou schemes. Results for different freestream Mach numbers and mesh topologies are discussed in order to assess the comparative performance of the various spatial discretization schemes. The flow is modeled by 3-D Euler equations through the use of a cell centered, face-based data structure finite voluma method applied in an unstructured grid context. Time integration of the system of equations is performed using an explicit, 5-stage, Runge-Kutta scheme. Mesh refinement routines are available in the original code and they are able to handle tetrahedra, hexahedra, triangular-base prisms and square-base pyramids. The full multigrid procedure is also available in the base code to accelerate the convergence to steady state. In the present work, the author has studied possible forms of integrating the multigrid and the mesh refinement procedures, which were both originally available in the base code. The results obtained provide evaluation and comparison of the present methods with regard to oblique shock wave capturing, as well as the behavior of property values such as pressure, density and Mach number contours. Finally, the work presents a discussion on the relative characteristics of each method.
23
Redding, Jeremy. "Deformation, Fragmentation and Vaporization of Volatile Liquid Droplets in Shock-Laden Environments." University of Cincinnati / OhioLINK, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1613745275535815.
24
Burym, A. S. "Development of the newest supersonicaircraft." Thesis, Київський національний університет технологій та дизайну, 2018. https://er.knutd.edu.ua/handle/123456789/11430.
25
Fojtl, Michal. "Výpočet aerodynamických charakteristik nosiče pro nízkou oběžnou dráhu." Master's thesis, Vysoké učení technické v Brně. Fakulta strojního inženýrství, 2017. http://www.nusl.cz/ntk/nusl-316914.
Abstract:
Master’s thesis deals with aerodynamic heating of launch vehicle during ascent phase by using CFD simulation. Ascent trajectory and payload fairing geometry is design using data of existing small launch vehicles. Critical flight regimes are identified using 2D calculations, and in these regimes analysis is performed by axially symmetric simulations. Simulation results are compared to values obtained from theoretical and semi-empirical calculations.
26
Crowell, Andrew R. "Model Reduction of Computational Aerothermodynamics for Multi-Discipline Analysis in High Speed Flows." The Ohio State University, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=osu1366204830.
27
Wallis, Scott Evan. "Innovative Transverse Jet Interaction Arrangements in Supersonic Crossflow." Thesis, Virginia Tech, 2001. http://hdl.handle.net/10919/36041.
Abstract:
The experiments on this project proceeded on the premise that adding an array of auxiliary jets behind a main jet injector would alleviate the large region of low pressure typically found downstream of a normal, sonic injector in supersonic flow and also possibly increase in intensity of the upstream high-pressure region. The secondary jet would, in theory, "push" the primary jet further into the flow, increasing the size of the obstacle as seen by the flow. The resulting increased high pressure upstream of the flow would increase the force on the body. Also, the presence of secondary jets would reduce the intensity of the primary jet's low-pressure region. These results would be beneficial to increase the force and decrease the nose-down moment associated with sonic, normal injection into a supersonic crossflow. Therefore, in application to hypersonic, high-altitude missile maneuvering, the firing of a thruster with such an array would result in both added force and a reduction of the moment usually associated with the pressure field on the missile. Such an array could allow the missile to perform purely translational maneuvers with less fuel, all the while keeping the target in view. To accomplish this task, some modern missiles use a second injector far downstream from the primary injector. This second injector's primary function is to negate the nose-down moment, and it adds little to the overall jet effectiveness.
To this end, two sets of experiments were conducted: one with low jet pressure ratio, Poj/P1 = 13.65, and low Mach number of 2.4 with Po,inf = 3.74 atm and To,inf = 293K for proof of concept and one at primary conditions Poj/P1 = 620, M1 = 4, Po,inf = 10.21 atm, To,inf = 293K. Spark Shadowgraphs were taken at both of these cases to study the structures present in the flow field and to qualitatively assess the effects of the secondary jet injectors. Placed under the Mach disk of the main jet, the secondary jets are hypothesized to push the plume of the main jet further up into the flow, increasing the force on the plate, and Shadowgraphs were used to test this hypothesis. Schlieren pictures were taken at the high M1, high-pressure ratio test case to further study the interaction of the secondary jets with the main jet. Pressure Sensitive Paint, PSP, was used in both cases to gain a greater understanding of the surface pressure near the injectors for different jet configurations. It was discovered that the addition of secondary jets could indeed both increase the force generated by the main jet and reduce the undesirable nose-down moment created by the main jet.
In the low M1, low pressure ratio conditions, the addition of one pair of jets manipulated the surface pressure such that the force on the plate increased by 17% and the nose-down moment was increased by 9% over the main jet only case. The further addition of one more pair of injectors increased the surface pressure force on the plate by 34% and increased the nose-down moment on the plate by 3% when compared to the Main Jet Only case. It is important to note that, these increases are due solely to the manipulation of the surface pressure force field and not the thrust of the secondary jets. The added thrust would increase the force on the plate and their position would insure an increase of a nose-up moment. One pair of secondary jets increases the injectant mass flow by about 2.3%. Therefore, the effects reported above are seen to be disproportionate to the amount of added injectant.
For the primary test conditions (M1 = 4, Poj/P1 = 620, Po = 10.21 atm, To = 293K) the addition of two pairs of secondary jets had a force increase of 62% and a nose-down moment decrease of 38% over that of the main jet only case. Three pairs increased the force 71% and the decreased the nose-down moment by 26% and four pairs increased the force 91% but increased the nose-down moment by 33%. These values do not account for the thrust of the secondary jets. Accounting for the beneficial effects of the thrust of the secondary jets, the force on the plate for two pairs of secondary jets increased the force 70% and decreased the moment 42%. Three pairs increased the force 83% and decreased the moment 35%. The increase of force for four pairs of secondary jets was 106% and the increase in nose-down moment was only 21%. A point of diminishing returns was reached. As more pairs of injectors are added further and further from the main injector, the beneficial force effects are offset by a growing moment penalty. By considering the locations of the secondary injectors to the main injector for both the low Mach number, low-pressure ratio tests and the main tests conditions, it can be surmised that the greatest benefit from the secondary jets can be extracted when the jets are placed within the main injector's downstream low-pressure region.Master of Science
28
Douay, Guy. "Modélisation et étude numérique de la turbulence compressible en écoulements supersoniques." Rouen, 1994. http://www.theses.fr/1994ROUES002.
Abstract:
Le sujet de la thèse porte sur les effets de la compressibilité forte dans les écoulements à très hautes vitesses (régimes supersoniques et hypersoniques). Le premier chapitre est consacré à l'élaboration des équations moyennes turbulentes et à la présentation du modèle de turbulence k-epsilon. Le deuxième chapitre développe la modélisation des termes dits de compressibilité qui comprennent: - la dissipation compressible EC, - la corrélation pression dilatation (fluctuations de pression/divergence des fluctuations de vitesse) PD. Une technique d'intégration numérique de type volumes finis, fondée sur l'algorithme de Mac Cormack est détaillee dans le troisième chapitre. L'exploitation des résultats numériques ainsi que leurs analyses constituent le dernier chapitre
29
Noubel, Hugo. "Etude expérimentale du comportement aérodynamique et optimisation des performances des planeurs hypersoniques dans des écoulements supersoniques et hypersoniques raréfiés." Electronic Thesis or Diss., Orléans, 2024. http://www.theses.fr/2024ORLE1001.
Abstract:
L’objectif principal de ce travail de thèse est de caractériser expérimentalement l’impact des effets visqueux sur les performances aérodynamiques des planeurs hypersoniques. Cette étude regroupe six écoulements basse pression (de 0,068 Pa à 71,11 Pa) de la soufflerie MARHy, quatre supersoniques (Mach 2 et Mach 4) et deux écoulements hypersoniques (Mach 20). Les maquettes expérimentées sont au nombre de 6 et l’objectif est d’étudier des géométries avec des degrés d’optimisation différents et de comprendre l’impact des effets visqueux sur chacune d’elles. Différents diagnostics ont été utilisés pour mener à bien cette étude : Tout d’abord, une balance aérodynamique a été développée pour pouvoir mesurer les forces de traînée et de portance des différentes configurations. Ensuite, les ondes de choc ont été visualisées à l’aide de visualisation par décharge luminescente. Enfin, une étude de pression pariétale a été menée sur deux planeurs hypersoniques. Ce travail de thèse permet d’établir une large base de données expérimentales sur les planeurs hypersoniques en régime raréfié. Les études de forces ont permis de quantifier l’évolution de la finesse au cours d’une rentrée atmosphérique en fonction du degré de raréfaction (paramètre de Tsien). Pour ce qui est des angles d’attaque, des formulations tenant compte des effets visqueux ont été établies et pourront être utilisées lors de l’optimisation des waveriders à hautes altitudes. Un planeur hypersonique tenant compte de ces données a été testé et est prometteur à haute altitude. Tout au long de cette thèse, les résultats mettent en évidence l’impact des effets visqueux sur les performances aérodynamiques des planeurs hypersoniquesThe main objective of this thesis work is to experimentally characterize the impact of viscous effects on the aerodynamic performance of hypersonic gliders. The study includes six low-pressure flows (ranging from 0.068 Pa to 71.11 Pa) from the MARHy wind tunnel, four supersonic flows (Mach 2 and Mach 4), and two hypersonic flows (Mach 20). A total of 6 models were tested, aiming to study geometries with different degrees of optimization and to understand the impact of viscous effects one ach of them. Various diagnostics were used to carry out this study : Firstly, an aerodynamic balance was developed to measure the drag and lift forces of the different configurations. Next, shock waves were visualized using glow discharge imaging. Finally, a parietal pressure study was carried out on two hypersonic gliders. This thesis work establishes a broad experimental database on hypersonic gliders in the rarefied regime. Force studies have enabled us to quantify the evolution of glide ratio during atmospheric re-entry as a function of the degree of rarefaction (Tsien parameter). As far as angles of attack are concerned, formulations taking viscous effects into account have been established and can be used in the optimization of high-altitude waveriders. A hypersonic glider incorporating these data has been tested and shows promise at high altitude. Throughout this thesis, results have highlighted the impact of viscous effects on the aerodynamic performance of hypersonic gliders
30
Nieberding, Zachary J. "An Investigation of Acoustic Wave Propagation in Mach 2 Flow." University of Cincinnati / OhioLINK, 2014. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1406881591.
31
Coumar, Sandra. "Etude des mécanismes physiques induits pas un actionneur plasma appliqué au contrôle d’écoulements raréfiés super/hypersoniques dans le cadre de rentrées atmosphériques." Thesis, Orléans, 2017. http://www.theses.fr/2017ORLE2025/document.
Abstract:
Ces dernières années, les missions spatiales bénéficient d'un regain d'intérêt. Cependant, lorsqu’arrive laphase d’entrée dans l’atmosphère, nous faisons encore face à d’importantes difficultés. Afin de répondre àce problème, une nouvelle technique est proposée : le contrôle par plasma pour augmenter la force detraînée sur le véhicule et ainsi, décroître sa vitesse. Dans cette thèse, un actionneur plasma est testé danstrois écoulements supersoniques (N1(M2-8Pa), N2(M4-8Pa) and N3(M4-71Pa)) et un hypersonique (M20-0.062Pa), ces écoulements étant simulés par la soufflerie MARHy.L’actionneur plasma induit des modifications de l’écoulement autour du modèle étudié, comme unemodification de la géométrie de l’onde de choc et une augmentation de l’angle de choc. Afin de mieuxcomprendre les phénomènes gouvernant ces modifications, la pression Pitot, la température surfacique etvolumique, les données électroniques et des mesures spectroscopiques ont été analysées. Les résultatsmontrèrent que deux types d’effets interviennent : thermiques (surface et volume) et l’ionisation. De plus, il aété démontré que ces effets n’ont pas la même importance suivant les conditions d’écoulements.L’actionneur plasma lui-même a été modifié dans un but d’amélioration. En particulier, deux types degénérateurs ont été étudiés pour alimenter la cathode : DC et pulsé. Finalement, il est montré que pour unepuissance de décharge de 80 W, une augmentation de 13% de la traînée et donc, une diminution de plus de25% des flux de chaleur peuvent être attendus. Par conséquent, les actionneurs plasma semblent être descandidats idéaux pour les missions spatiales et les (r)entrées atmosphériqueSpace missions are arousing renewed interest in these recent years. However, when coming to the entryinto the atmosphere, major issues are still to be considered. To answer this problem, a new Entry DescentLanding technique is proposed: plasma actuation to increase the drag force over the vehicle body and thus,decrease its speed. In this thesis, a plasma actuator is tested in three supersonic rarefied flows (N1(M2-8Pa), N2(M4-8Pa) and N3(M4-71Pa)) and a hypersonic one (M20-0.062Pa), all generated by the wind tunnelMARHy.The plasma actuator induces flow modifications over the studied model, such as a change in the shock waveshape and an increase in the shock wave angle. In order to better understand the phenomena governingthese modifications, Pitot pressure, surface and gas temperature, electron data and spectroscopicmeasurements were analyzed. The results shown that two types of effects are involved: thermal (bulk andsurface) and ionization. Moreover, it was demonstrated that these effects had not the same importancedepending on the flow conditions.The plasma actuator was also modified in order to improve it. In particular, two types of generators wereused to biase the cathode: DC and pulsed. Finally, it was shown that, for a discharge power of 80 W, a 13%increase in the drag force could be expected and thus, a decrease in the heat load over the model body ofmore than 25%. Therefore, plasma actuators seem to be promising applications for space missions andatmospheric entries
32
André, Thierry. "Contrôle actif de la transition laminaire-turbulent en écoulement hypersonique." Thesis, Orléans, 2016. http://www.theses.fr/2016ORLE2022/document.
Abstract:
Lors d’un vol hypersonique (Mach 6, 20 km d’altitude) la couche limite se développant sur l’avant-corps d’un véhicule hypersonique est laminaire. Cet état cause un désamorçage du moteur (statoréacteur) assurant la propulsion du véhicule. Pour pallier ce problème, il faut forcer la transition de la couche limite á l’aide d’un dispositif de contrôle dont l’effet est permanent (passif) ou modulable (actif) pendant le vol. Dans ce travail, nous analysons l’efficacité d’un dispositif actif d’injection d’air á la paroi pour forcer la transition de la couche limite sur un avant-corps générique. L’interaction jet d’air/couche limite est simulée numériquement avec une approche aux grandes échelles (LES). Une étude paramétrique sur la pression d’injection permet de quantifier l’efficacité du jet á déstabiliser la couche limite. L’influence des conditions de vol (altitude, Mach) sur la transition est également étudiée. Une analyse des résultats de simulation par Décomposition en Modes Dynamiques (DMD) est menée pour comprendre quels sont les modes dynamiques responsables de la transition et les mécanismes sous-jacents. Des essais dans la soufflerie silencieuse de l’université de Purdue (BAM6QT) ont été effectués pour tester expérimentalement l’efficacité des dispositifs passifs (rugosité isolée en forme de losange) et actifs (mono-injection d’air) pour faire transitionner la couche limite. Une peinture thermo-sensible et des capteurs de pression (PCB, Kulite) ont été utilisés pour déterminer la nature de la couche limite. Les résultats de ce travail montrent qu’une injection sonique suffit pour forcer la couche limite. On observe des essais, que pour une même hauteur de pénétration, les rugosités isolées sont moins efficaces que les jets (mono injection) pour déstabiliser la couche limiteDuring a hypersonic flight (Mach 6, 20 km altitude), the boundary layer developing on the forebody of a vehicle is laminar. This state may destabilize the scramjet engine propelling the vehicle. To overcome this problem during the flight, the boundary layer transition has to be forced using a control device whose effect is fixed (passive) or adjustable (active). In this work, we analyze the efficiency of a jet in crossflow in forcing the boundary layer transition on a generic forebody. The flow is computed with a Large Eddy Simulations (LES) approach. A parametric study of the injection pressure allows the efficiency of the jet in tripping the boundary layer to be quantified. The influence of flight conditions (Mach, altitude) on the transition is also studied. Dynamic Mode Decomposition (DMD) is applied to the simulation results to determine the transition leading to dynamic modes and to understand underlying transition mechanisms. Experiments in the Purdue University quiet wind tunnel (BAM6QT) were performed to quantify the efficiency of a passive transition device (diamond roughnesses) and an active transition device (single air jet) in tripping the boundary layer. A thermo-sensitive paint and pressure transducers (Kulite, PCB) were used to determine the state of the boundary layer on the generic forebody. Experimental and numerical results show a sonic injection is sufficient to induce transition. We observe from the experiments that for the same penetration height, a single roughness is less efficient than a single air jet in destabilizing the boundary layer
33
劉台安. "On waverider shapes applied to future supersonic and hypersonic configuration." Thesis, 1990. http://ndltd.ncl.edu.tw/handle/43529985420926245312.
34
Munuswamy, Nithiyaraj. "Jet Injection into Supersonic Crossflow: Flowfield and Mixing studies." Thesis, 2018. https://etd.iisc.ac.in/handle/2005/4738.
Abstract:
Supersonic air-breathing engines or scramjets are perhaps the most important technological
hurdles towards the development of hypersonic transportation vehicles for both easy
access to space and for missile related applications. In these engines, the air entering the
combustor must remain supersonic, which significantly brings down the time available within
the engine for fuel-air mixing and combustion. There are many possible strategies for fuel
injection in such engines; the transverse jet injection into supersonic crossflow being one.
The principle parameter in jet in crossflow studies is the momentum flux ratio (J), which
is defined as the momentum flux ratio between the jet and the cross-flow. Compared to the
subsonic crossflow case, the supersonic crossflow has additional complexities due to the
presence of different shock structures in addition to the many different vortical structures.
The aim of the present work is to experimentally study the flowfield and mixing associated
with sonic jet injection into a supersonic crossflow. In addition to detailed flowfield
and mixing studies of the basic steady jet configuration, enhancement of mixing using a
newly developed passively modulated (injection) jet has also been studied. In the latter case,
a large range of jet modulation frequencies ( f ) characterized by the ratio of the modulation
frequency to the natural shear layer frequency ( f / fsl) have been investigated ( f / fsl = 0.12
to 1.31). In all cases, the flowfield is investigated using Particle Image Velocimetry (PIV)
to characterize the different flow features and the penetration of the jet into the crossflow.
Mixing studies have also been carried out using acetone Planar Laser Induced Fluorescence
(PLIF), with the injected jet containing acetone vapours, which is then tracked after injection
into the crossflow to quantify mixing.
In summary, the modulated jet shows significant changes in jet penetration, unsteady
motions of the jet, and the mixing of the jet into the crossflow. Among the large number of
cases studied, the penetration and the mixing are found to be higher at modulation frequencies
corresponding to the sub-harmonics of the shear layer frequency. The maximum penetration
is seen close to f / fsl = 0.26, but the maximum mixedness of the injected jet is seen to occur
at f / fsl = 0.46, where the penetration is higher than the base case, but not the highest
35
Joarder, Ratan. "Demonstration Of Supersonic Combustion In A Combustion Driven Shock-Tunnel." Thesis, 2009. https://etd.iisc.ac.in/handle/2005/1005.
Abstract:
For flights beyond Mach 6 ramjets are inefficient engines due to huge total pressure loss in the normal shock systems, combustion conditions that lose a large fraction of the available chemical energy due to dissociation and high structural loads. However if the flow remains supersonic inside the combustion chamber, the above problems could be alleviated and here the concept of SCRAMJET(supersonic combustion ramjet) comes into existence. The scramjets could reduce launching cost of satellites by carrying only fuel and ingesting oxygen from atmospheric air. Further applications could involve defense and transcontinental hypersonic transport.
In the current study an effort is made to achieve supersonic combustion in a ground based short duration test facility(combustion driven shock-tunnel), which in addition to flight Mach number can simulate flight Reynolds number as well. In this study a simple method of injection i.e. wall injection of the fuel into the combustion chamber is used. The work starts with threedimensional numerical simulation of a non-reacting gas(air) injection into a hypersonic cross-flow of air to determine the conditions in which air penetrates reasonably well into the cross-flow. Care is taken so that the process does not induce huge pressure loss due to the bow shock which appears in front of the jet column. The code is developed in-house and parallelized using OpenMp model. This is followed by experiments on air injection into a hypersonic cross-flow of air in a conventional shock-tunnel HST2 existing in IISc. The most tricky part is synchronization of injection with start of test-flow in such a short duration(test time 1 millisecond) facility.
Next part focuses on numerical simulations to determine the free-stream conditions, mainly the temperature and pressure of air, so that combustion takes place when hydrogen is injected into a supersonic cross-flow of air. The simulations are two-dimensional and includes species conservation equations and source terms due to chemical reactions in addition to the Navier-Stokes equations. This code is also built in-house and parallelized because of more number of operations with the inclusion of species conservation equations and chemical non-equilibrium. However, the predicted conditions were not achievable by HST2 due to low stagnation conditions of HST2.
Therefore, a new shock-tunnel which could produce the required conditions is built. The new tunnel is a combustion driven shock-tunnel in which the driver gas is at higher temperature than conventional shock-tunnel. The driver gas is basically a mixture of hydrogen, oxygen and helium at a mole ratio of 2:1:10 initially. The mixture is ignited by spark plugs and the hydrogen and oxygen reacts releasing heat. The heat released raises the temperature of the mixture which is now predominantly helium and small fractions of water vapour and some radicals. The composition of the driver gas and initial pressure are determined through numerical simulations.
Experiments follow in the new tunnel on hydrogen injection into a region of supersonic cross-flow between two parallel plates with a wedge attached to the bottom plate. The wedge reduces the hypersonic free-stream to Mach 2. A high-speed camera monitors the flow domain around injection point and sharp rise in luminosity is observed. To ascertain whether the luminosity is due to combustion or not, two more driven gases namely nitrogen and oxygen-rich air are used and the luminosity is compared. In the first case, the free-stream contains no oxygen and luminosity is not observed whereas in the second case higher luminosity than air driver case is visible. Additionally heat-transfer rates are measured at the downstream end of the model and at a height midway between the plates. Similar trend is observed in the relative heat-transfer rates. Wall static pressure at a location downstream of injection port is also measured and compared with numerical simulations. Results of numerical simulations which are carried out at the same conditions as of experiments confirm combustion at supersonic speed.
Experiments and numerical simulations show presence of supersonic combustion in the setup. However, further study is necessary to optimize the parameters so that thrust force could be generated efficiently.
36
Joarder, Ratan. "Demonstration Of Supersonic Combustion In A Combustion Driven Shock-Tunnel." Thesis, 2009. http://hdl.handle.net/2005/1005.
Abstract:
For flights beyond Mach 6 ramjets are inefficient engines due to huge total pressure loss in the normal shock systems, combustion conditions that lose a large fraction of the available chemical energy due to dissociation and high structural loads. However if the flow remains supersonic inside the combustion chamber, the above problems could be alleviated and here the concept of SCRAMJET(supersonic combustion ramjet) comes into existence. The scramjets could reduce launching cost of satellites by carrying only fuel and ingesting oxygen from atmospheric air. Further applications could involve defense and transcontinental hypersonic transport.
In the current study an effort is made to achieve supersonic combustion in a ground based short duration test facility(combustion driven shock-tunnel), which in addition to flight Mach number can simulate flight Reynolds number as well. In this study a simple method of injection i.e. wall injection of the fuel into the combustion chamber is used. The work starts with threedimensional numerical simulation of a non-reacting gas(air) injection into a hypersonic cross-flow of air to determine the conditions in which air penetrates reasonably well into the cross-flow. Care is taken so that the process does not induce huge pressure loss due to the bow shock which appears in front of the jet column. The code is developed in-house and parallelized using OpenMp model. This is followed by experiments on air injection into a hypersonic cross-flow of air in a conventional shock-tunnel HST2 existing in IISc. The most tricky part is synchronization of injection with start of test-flow in such a short duration(test time 1 millisecond) facility.
Next part focuses on numerical simulations to determine the free-stream conditions, mainly the temperature and pressure of air, so that combustion takes place when hydrogen is injected into a supersonic cross-flow of air. The simulations are two-dimensional and includes species conservation equations and source terms due to chemical reactions in addition to the Navier-Stokes equations. This code is also built in-house and parallelized because of more number of operations with the inclusion of species conservation equations and chemical non-equilibrium. However, the predicted conditions were not achievable by HST2 due to low stagnation conditions of HST2.
Therefore, a new shock-tunnel which could produce the required conditions is built. The new tunnel is a combustion driven shock-tunnel in which the driver gas is at higher temperature than conventional shock-tunnel. The driver gas is basically a mixture of hydrogen, oxygen and helium at a mole ratio of 2:1:10 initially. The mixture is ignited by spark plugs and the hydrogen and oxygen reacts releasing heat. The heat released raises the temperature of the mixture which is now predominantly helium and small fractions of water vapour and some radicals. The composition of the driver gas and initial pressure are determined through numerical simulations.
Experiments follow in the new tunnel on hydrogen injection into a region of supersonic cross-flow between two parallel plates with a wedge attached to the bottom plate. The wedge reduces the hypersonic free-stream to Mach 2. A high-speed camera monitors the flow domain around injection point and sharp rise in luminosity is observed. To ascertain whether the luminosity is due to combustion or not, two more driven gases namely nitrogen and oxygen-rich air are used and the luminosity is compared. In the first case, the free-stream contains no oxygen and luminosity is not observed whereas in the second case higher luminosity than air driver case is visible. Additionally heat-transfer rates are measured at the downstream end of the model and at a height midway between the plates. Similar trend is observed in the relative heat-transfer rates. Wall static pressure at a location downstream of injection port is also measured and compared with numerical simulations. Results of numerical simulations which are carried out at the same conditions as of experiments confirm combustion at supersonic speed.
Experiments and numerical simulations show presence of supersonic combustion in the setup. However, further study is necessary to optimize the parameters so that thrust force could be generated efficiently.
37
Zhikharev, Constantin N. "Interaction theory for hypersonic separation and supersonic flow past a flexible wall /." Diss., 1999. http://gateway.proquest.com/openurl?url_ver=Z39.88-2004&rft_val_fmt=info:ofi/fmt:kev:mtx:dissertation&res_dat=xri:pqdiss&rft_dat=xri:pqdiss:9935191.
38
Hou, Chang-Huan, and 侯昌桓. "A Study of Supersonic, Transonic, and Hypersonic Small Perturbation Theory and Their Verifications." Thesis, 1999. http://ndltd.ncl.edu.tw/handle/20199326053444391078.
Abstract:
碩士淡江大學
航空太空工程學系
87
The essential point of view in this paper is to re-examine the small perturbation theory of supersonic, transonic, and hypersonic flow. The asymptotic solutions of boundary value problems of centered expansion and compression waves for supersonic, transonic, and hypersonic nonlinear small disturbance equations are exhibited. The asymptotic solution of compression wave is verified by the exact solution of the oblique shock wave theory. Due to the singularity at the corner, the supersonic linearized theory is invalid for calculating the flow field of corner flow. By using the method of strained coordinates, the nonlinear asymptotic theory of supersonic flow is built and it gives the correct solution for corner flow. The shock wave position and the pressure coefficient on the wedge for supersonic, transonic, and hypersonic small disturbance theory are given to the first order; and the results agree well with first order asymptotic solution of the exact solution of the oblique shock wave theory. Keywords : supersonic, transonic, and hypersonic nonlinear small disturbance theory, corner flow, oblique shock wave theory
39
Chen, Yang Shie, and 楊士震. "The Study of Internal and External Compressible Viscous Flow for Supersonic and Hypersonic." Thesis, 1993. http://ndltd.ncl.edu.tw/handle/11663006880731778965.
40
(6632393), Ian Avalon Hall. "Simulating Scramjet Behavior: Unstart Prediction in a Supersonic, Turbulent Inlet-Isolator Duct Flow." Thesis, 2019.
Abstract:
In the pursuit of developing hypersonic cruise vehicles, unstart is a major roadblock to achieving stable flight. Unstart occurs when a sudden instability in the combustor of a vehicle’s propulsion system creates an instantaneous pressure rise that initiates a shock. This shock travels upstream out of the inlet of the vehicle, until it is ejected from the inlet and creates a standing shockwave that chokes the flow entering the vehicle, thereby greatly reducing its propulsive capability. In severe cases, this can lead to the loss of the vehicle. This thesis presents the results of a computational study of the dynamics of unstart near Mach 5 and presents some possible precursor signals that may indicate its presence in flight. Using SU2, an open-source CFD code developed at Stanford University, the Unsteady Reynolds-Averaged Navier-Stokes equations are used to develop a model for flow in a scramjet inlet-isolator geometry, both in the fully started state and during unstart. The results of these calculations were compared against experimental data collected by J. Wagner, at the University of Texas, Austin. In the present computations, unstart was initiated through the use of an artificial body force, which mimicked a moveable flap used in the experiments. Once the results of the code were validated against these experiments, a selection of parametric studies were conducted to determine how the design of the inlet-isolator by Wagner affected the flow, and thus how generalizable the results can be. In addition, precursor signals indicative of unstart were identified for further study and examined in the different parametric studies. It was found that a thick boundary layer is conducive to a stronger precursor signal and a slower unstart. In addition, an aspect ratio closer to 1:1 promotes flow mixing and reduces the unstart speed and strength. Moreover, an aspect ratio in this range reduces the precursor signal strength but, if a thick boundary layer is present, will smear the signal out over a larger area, potentially making it easier to detect.
41
Lee, Sung-Wei, and 李嵩蔚. "A Study of Nonlinear Theories of Supersonic, Transonic, and Hypersonic Flow in Far Field and Whole Field." Thesis, 2002. http://ndltd.ncl.edu.tw/handle/32102230248654132219.
Abstract:
碩士淡江大學
航空太空工程學系
90
The essential point of view in this paper is to re-examine the small perturbation theories of supersonic, transonic, and hypersonic flow in the far field and whole field. Due to the singularity at the corner, the supersonic linearized theory is invalid for calculating the flow field of corner flow. By using the method of strained coordinates, the nonlinear theory of supersonic flow is built and it gives the correct solution for the whole field of corner flow. By using the asymptotic expansion, a systematic and rigorous approach for obtaining the supersonic nonlinear far field equation ( inviscid Burgers equation ) is exhibited. The transonic and hypersonic far field equations are derived from the supersonic far field equation by considering a general expansion of free stream Mach number in terms of transonic and hypersonic similarity parameters together with stretched coordinates. The asymptotic solution of the boundary-value problem of corner flow are obtained from the nonlinear theories of supersonic, transonic, and hypersonic flow in far field and whole field. The far field and whole field results of supersonic, transonic, and hypersonic corner flow are also verified in the physical coordinates and far field coordinates respectively. It is worth mentioning that the mathematical structures of the far field and whole field equations of the supersonic flow are similar in the physical coordinates, but they are quite different in the far field coordinates. The problem of the supersonic, transonic, and hypersonic flow past a parabolic-shaped airfoil are solved in the far field including the similar structure of leading and trailing curved shocks.
42
Devaraj, Manoj Kumar K. "Physical insights into unstart dynamics of a hypersonic mixed compression intake." Thesis, 2021. https://etd.iisc.ac.in/handle/2005/5655.
Abstract:
Hypersonic air-breathing cruise vehicles powered by supersonic combustion ramjet engines
are the potential candidate for future space and defense applications. The air intake of the
scramjet engine is a vital component that uses shock waves to compress the air to pressure
and temperatures suitable for supersonic combustion. Understanding the unstart dynamics of
such intakes is of prime importance for the seamless operation of scramjet intakes. While the
unstart dynamics in supersonic intakes are studied widely by various researchers, only a few
such studies are reported in hypersonic intakes. The mechanisms associated with the same are
not clearly understood. In the current work, a design optimization framework is established
by coupling (a) oblique-shock theory and Non-dominated Sorting Genetic Algorithm II
(NSGA-II) and (b) Computational fluid dynamics (CFD) and NSGA - II to minimize total
pressure loss and maximize intake exit temperature of planar mixed compression intake at
a design Mach number of 6. The ramp and cowl angles constitute the design space. The
intake with maximum exit temperature is chosen to study its unstart dynamics using a
combination of experiments in a hypersonic wind tunnel (M = 6 and Re = 8.86 × 106/m)
and unsteady numerical investigations using the open-source suite SU2. The intake model
is equipped with a movable cowl and flap to study the internal contraction and throttling
induced unstart. Simultaneous pressure measurements and schlieren flow visualization are
carried out to study unsteady flow physics associated with intake unstart. The dynamic
content in the flow is analyzed using Fast Fourier Transform (FFT) and spectrogram of the
unsteady pressure signal and Dynamic Mode Decomposition (DMD) of the schlieren images
and density contours.
In this work, two different modes of shock oscillation during unstart are observed when
the flap is moved while the cowl is held stationary. At ICR = 1.19, the intake shows started
behavior for throttling ratio up to 0.31, and a dual behavior, where it remains started in
dynamic flap runs but unstarted in fixed flap runs for throttling ratios of 0.35 and 0.42. The
intake exhibits a staged evolution to a large amplitude oscillatory unstart for throttling ratios
of 0.55 and 0.69, with frequencies of 950 and 1100 Hz, respectively. A staged evolution (5
stages) to a subsonic spillage oscillatory unstart is detailed using corroborative evidence from
both time-resolved schlieren and pressure measurements. The ramp side separation bubble
drives the high amplitude oscillatory unstart. At ICR = 1.37, the shear layer emanating
from the triple point of shock interaction drives the low amplitude oscillatory unstart with a
dominant frequency of about 3.7 kHz for a throttling ratio of 0.69. A criterion for demarcating
the modes of unstart is evolved using current and previous data. The actual shock on lip
condition during started operation demarcates the two modes of oscillatory unstart. Unsteady
numerical computations are performed to study the effect of enthalpy on the unstart frequency.
The frequency of unstart varies linearly with stagnation acoustic speed and is an appropriate
velocity scale. During unstart, the extent of the subsonic region is the appropriate length
scale to be used in the quarter-wave resonance model to estimate unstart frequency pertaining
to high mechanical blockage
43
Tichenor, Nathan R. "Characterization of the Influence of a Favorable Pressure Gradient on the Basic Structure of a Mach 5.0 High Reynolds Number Supersonic Turbulent Boundary Layer." 2010. http://hdl.handle.net/1969.1/ETD-TAMU-2010-08-8407.
Abstract:
High-speed high Reynolds number boundary layer flows with mechanical non-equilibrium effects have numerous practical applications; examples include access-to-space ascent, re-entry and descent, and military hypersonic systems. However, many of the basic turbulent flow processes in this regime are poorly understood and are beyond the realm of modern direct numerical simulations Previous studies have shown that curvature driven pressure gradients significantly alter the state of the turbulence in high-speed boundary layers; the turbulence levels have been shown to decrease by large amounts (up to 100 percent) and the Reynolds shear stress has been shown to change sign. However, most of our understanding is based on point measurement techniques such as hot-wire and Laser Doppler anemometry acquired at low to moderate supersonic Mach numbers (i.e., M = 2-3). After reviewing the available literature, the following scientific questions remain unanswered pertaining to the effect of favorable pressure gradients:
(1) How is state of the mean flow and turbulence statistics altered?
(2) How is the structure of wall turbulence; break-up, stretch or a combination?
(3) How are the Reynolds stress component production mechanisms altered?
(4) What is the effect of Mach number on the above processes?
To answer these questions and to enhance the current database, an experimental analysis was performed to provide high fidelity documentation of the mean and turbulent flow properties using two-dimensional particle image velocimetry (PIV) along with flow visualizations of a high speed (M4.88=), high Reynolds number (Re36,000θ≈) supersonic turbulent boundary layer with curvature-driven favorable pressure gradients (a nominally zero, a weak, and a strong favorable pressure gradient). From these data, detailed turbulence analyses were performed including calculating classical mean flow and turbulence statistics, examining turbulent stress production, and performing quadrant decomposition of the Reynolds stress for each pressure gradient case.
It was shown that the effect of curvature-driven favorable pressure gradients on the turbulent structure of a supersonic boundary layer was significant. For the strong pressure gradient model, the turbulent shear stress changed sign throughout the entire boundary layer; a phenomena was not observed to this magnitude in previous studies. Additionally, significant changes were seen in the turbulent structure of the boundary layer. It is believed that hairpin vortices organized within the boundary layer are stretched and then broken up over the favorable pressure gradient. Energy from these hairpin structures is transferred to smaller turbulent eddies as well as back into the mean flow creating a fuller mean velocity profile. It was determined that the effects of favorable pressure gradients on the basic structure of a turbulent Mach 5.0 boundary layer were significant, therefore increasing the complexity of computational modeling.
44
Griffiths, Alan David. "Development and demonstration of a diode laser sensor for a scramjet combustor." Phd thesis, 2005. http://hdl.handle.net/1885/47106.
Abstract:
Hypersonic vehicles, based on scramjet engines, have the potential to deliver
inexpensive access to space when compared with rocket propulsion. The technology,
however, is in its infancy and there is still much to be learned from
fundamental studies.
¶
Flows that represent the conditions inside a scramjet engine can be generated in ground tests using a free-piston shock tunnel and a combustor model. These facilities provide a convenient location for fundamental studies and principles learned during ground tests can be applied to the design of a full-scale vehicle.¶ A wide range of diagnostics have been used for studying scramjet flows, including surface measurements and optical visualisation techniques.¶ The aim of this work is to test the effectiveness of tunable diode laser absorption spectroscopy (TDLAS) as a scramjet diagnostic.¶ ...
45
Mahapatra, Debabrata. "Investigation Of Ramp/Cowl Shock Interaction Processes Near A Generic Scramjet Inlet At Hypersonic Mach Number." Thesis, 2008. https://etd.iisc.ac.in/handle/2005/807.
Abstract:
One of the major technological innovations that are necessary for faster and cheaper access-to-space will be the commercial realization of supersonic combustion jet engines (SCRAMJET). The establishment of the flow through the inlet is one the prime requirement for the success of a SCRAMJET engine. The flow through a SCRAMJET inlet is dominated by inviscid /viscous coupling, transition, shock-shock interaction, shock boundary layer interaction, blunt leading edge effects and flow profile effects. Although the literature is exhaustive on various aspects of flow features associated with SCRAMJET engines, very little is known on the fundamental gasdynamic features dictating the flow establishment in the SCRAMJET inlet. On one hand we need the reduction of flight Mach number to manageable supersonic values inside the SCRAMJET combustor, but on the other hand we have to achieve this with minimum total pressure loss. Hence the dynamics of ramp/cowl shock interaction process ahead of the inlet has a direct bearing on the quality and type of flow inside the SCRAMJET engine. There is virtually no data base in the open literature focusing specifically on the cowl/ramp shock interactions at hypersonic Mach numbers.
Hence in this backdrop, the main aim of the present investigation is to systematically understand the ramp/cowl shock interaction processes in front of a generic inlet model. Since we are primarily concerned with the shock interaction process ahead of the cowl all the investigations are carried out without any fuel injection. Variable geometry is necessary if we want to operate the inlet for a wide range of Mach numbers in actual flight. The investigation mainly comprises of three variable geometry configurations; namely, variation of contraction ratios at 00 cowl (CR 8.4, 5.0 and 4.3), variation of cowl length for a given chamber height (four lengths of cowls at 10 mm chamber height) and variation of cowl angle (three angles cowl each for two chamber heights). The change in cowl configuration results in different ramp/cowl shock interaction processes affecting the performance of the inlet. Experiments are performed in IISc hypersonic shock tunnel HST 2 (test time ~ 1 ms) at two nominal Mach numbers 8.0 and 5.74 for design and off-design testing conditions. Exhaustive numerical simulations are also performed to compliment the experiments. Further the effect of concentrated energy deposition on forebody /cowl shock interactions has also been investigated.
A 2D, planar, single ramp scramjet inlet model has been designed and fabricated along with various cowl geometries and tested in a hypersonic shock tunnel to characterize the forebody/cowl shock interaction process for different inlet configurations. Further a DC plasma power unit and a plasma torch have been designed, developed and fabricated to serve as energy source for conducting flow-alteration experiments in the inlet model. The V-I characteristics of the plasma torch is studied and an estimation of plasma temperature is also performed as a part of characterizing the plasma flame. Initial standardization experiments of blunt body flow field alteration using the plasma torch and hence its drag reduction, are performed to check the torch’s suitability to be used as a flow-altering device in a shock tunnel. The plasma torch is integrated successfully with the inlet model in a shock tunnel to perform experiments with plasma jet as the energy source. The above experiments are first of its kind to be conducted in a shock tunnel. They are performed at various pressure ratios and supply currents. Time resolved schlieren flow visualization using Phantom 7.1 (Ms Vision Research USA) high speed camera, surface static pressure measurements inside a generic inlet using miniature kulite transducer and surface convective heat transfer rate measurements inside a generic inlet using platinum thin film sensors deposited on Macor substrate are some of the shock tunnel flow diagnostics that have been used in this study.
Some of the important conclusions from the study are:
• Experiments performed at different contraction ratios show different shock patterns. At CR 8.4, the SOL condition is satisfied, but the flow gets choked due to over contraction and flow through inlet is not established. For CR 5.0, formation of a small Mach stem before the chamber is observed with the reflection point on the cowl and the weak reflected shock entering inside the chamber. The Mach stem grows with time. For CR 4.3, the forebody/cowl shock interference created an Edney’s Type II shock interaction pattern. However, at off-design conditions, for CR 5 the shock reflection is regular and at CR 4.3, the Edney’s Type II pattern lasts for a short time.
• For all lengths of cowl tested, 131mm and 141mm showed Edney’s Type II shock interference where as 151mm showed Edney’s Type I pattern at design condition. In all cases the flow is choked for high contraction ratio. At off-design condition these shock patterns do not last for the entire test time but rather it becomes a lambda pattern with the normal shock before the inlet.
• For inlet configurations with cowl angle other than 00, the flow is found to be established for all cases at designed condition and except for 100 cowl at off-design condition.
• For CR 8.4 the peak value of pressure (~1.7x104 Pa) occurs at a location of 151mm, where as for CR 5.0 and 4.3 they occur at 188mm and 206mm having values ~1.6x104 Pa and ~1.4x104 Pa respectively. These locations indicate the likely locations of shock impingements inside the chamber.
• For cowl angle of 00 for a 10 mm chamber the maximum pressure value recorded is ~1.7x104 Pa whereas for 100 and 200 cowl it is ~1.1x104 Pa and 1.2 x104 Pa respectively. This is because in the first case the inlet is choked because of over contraction whereas in the other two cases the CR is less and flow is established inside the inlet.
• The average heat transfer rates of last four heat transfer gauges (180 mm, 190 mm, 200 mm and 210 mm from the forebody tip) for all lengths of cowls tested are found to be almost same (~ 20 W/cm2). This is because the flow is choked in all these cases. The numerical simulation also shows uniform distribution here, consistent with the experimental findings.
• The locations of heat transfer peaks for 100 cowl at design condition can be observed to be occurring at 170 mm and 200 mm from the forebody tip having values ~44 W/cm2 and ~39 W/cm2 respectively. For a 200 cowl they seem to be occurring at 170 mm and 180 mm from the forebody tip having values ~50 W/cm2 and ~30 W/cm2. These locations indicate the likely locations of shock impingements inside the chamber. With the evolution of concept of upstream fuel injection in recent times these may the most appropriate locations for fuel injection.
• At higher jet pressure ratios the plasma jet/ramp shock interaction results in a lambda shock pattern with the triple point forming vertically above the cowl level. This means the normal shock stands in front of the inlet making a part of the flow entering the inlet subsonic. The reflected shock from the triple point also separates the ramp boundary layer.
• At lower jet pressure ratios the triple point is formed below the cowl level and the flow entering inside the inlet is supersonic. The reflected shock interacts with the cowl shock and a weak separation shock is seen.
• Experiments are performed with concentrated DC electric discharge as energy source. Even though the amount of energy dumped here is less than 0.25% of the total energy it creates a perceptible disturbance in the flow.
• Experiments are also performed to see the effect of electric discharge as energy source on height of Mach stem for a given inlet configuration. Deposition of energy in the present location does not seem to alter the Mach stem height.
However more experiments need to be performed by varying the energy location to see its effect. Non-intrusive energy sources like microwave and lasers can be thought of as options for depositing energy to study its effect on Mach stem height. Since they provide more flexibility on varying the location of energy the optimum location of energy can be found out for highest reduction of Mach stem height.
46
Mahapatra, Debabrata. "Investigation Of Ramp/Cowl Shock Interaction Processes Near A Generic Scramjet Inlet At Hypersonic Mach Number." Thesis, 2008. http://hdl.handle.net/2005/807.
Abstract:
One of the major technological innovations that are necessary for faster and cheaper access-to-space will be the commercial realization of supersonic combustion jet engines (SCRAMJET). The establishment of the flow through the inlet is one the prime requirement for the success of a SCRAMJET engine. The flow through a SCRAMJET inlet is dominated by inviscid /viscous coupling, transition, shock-shock interaction, shock boundary layer interaction, blunt leading edge effects and flow profile effects. Although the literature is exhaustive on various aspects of flow features associated with SCRAMJET engines, very little is known on the fundamental gasdynamic features dictating the flow establishment in the SCRAMJET inlet. On one hand we need the reduction of flight Mach number to manageable supersonic values inside the SCRAMJET combustor, but on the other hand we have to achieve this with minimum total pressure loss. Hence the dynamics of ramp/cowl shock interaction process ahead of the inlet has a direct bearing on the quality and type of flow inside the SCRAMJET engine. There is virtually no data base in the open literature focusing specifically on the cowl/ramp shock interactions at hypersonic Mach numbers.
Hence in this backdrop, the main aim of the present investigation is to systematically understand the ramp/cowl shock interaction processes in front of a generic inlet model. Since we are primarily concerned with the shock interaction process ahead of the cowl all the investigations are carried out without any fuel injection. Variable geometry is necessary if we want to operate the inlet for a wide range of Mach numbers in actual flight. The investigation mainly comprises of three variable geometry configurations; namely, variation of contraction ratios at 00 cowl (CR 8.4, 5.0 and 4.3), variation of cowl length for a given chamber height (four lengths of cowls at 10 mm chamber height) and variation of cowl angle (three angles cowl each for two chamber heights). The change in cowl configuration results in different ramp/cowl shock interaction processes affecting the performance of the inlet. Experiments are performed in IISc hypersonic shock tunnel HST 2 (test time ~ 1 ms) at two nominal Mach numbers 8.0 and 5.74 for design and off-design testing conditions. Exhaustive numerical simulations are also performed to compliment the experiments. Further the effect of concentrated energy deposition on forebody /cowl shock interactions has also been investigated.
A 2D, planar, single ramp scramjet inlet model has been designed and fabricated along with various cowl geometries and tested in a hypersonic shock tunnel to characterize the forebody/cowl shock interaction process for different inlet configurations. Further a DC plasma power unit and a plasma torch have been designed, developed and fabricated to serve as energy source for conducting flow-alteration experiments in the inlet model. The V-I characteristics of the plasma torch is studied and an estimation of plasma temperature is also performed as a part of characterizing the plasma flame. Initial standardization experiments of blunt body flow field alteration using the plasma torch and hence its drag reduction, are performed to check the torch’s suitability to be used as a flow-altering device in a shock tunnel. The plasma torch is integrated successfully with the inlet model in a shock tunnel to perform experiments with plasma jet as the energy source. The above experiments are first of its kind to be conducted in a shock tunnel. They are performed at various pressure ratios and supply currents. Time resolved schlieren flow visualization using Phantom 7.1 (Ms Vision Research USA) high speed camera, surface static pressure measurements inside a generic inlet using miniature kulite transducer and surface convective heat transfer rate measurements inside a generic inlet using platinum thin film sensors deposited on Macor substrate are some of the shock tunnel flow diagnostics that have been used in this study.
Some of the important conclusions from the study are:
• Experiments performed at different contraction ratios show different shock patterns. At CR 8.4, the SOL condition is satisfied, but the flow gets choked due to over contraction and flow through inlet is not established. For CR 5.0, formation of a small Mach stem before the chamber is observed with the reflection point on the cowl and the weak reflected shock entering inside the chamber. The Mach stem grows with time. For CR 4.3, the forebody/cowl shock interference created an Edney’s Type II shock interaction pattern. However, at off-design conditions, for CR 5 the shock reflection is regular and at CR 4.3, the Edney’s Type II pattern lasts for a short time.
• For all lengths of cowl tested, 131mm and 141mm showed Edney’s Type II shock interference where as 151mm showed Edney’s Type I pattern at design condition. In all cases the flow is choked for high contraction ratio. At off-design condition these shock patterns do not last for the entire test time but rather it becomes a lambda pattern with the normal shock before the inlet.
• For inlet configurations with cowl angle other than 00, the flow is found to be established for all cases at designed condition and except for 100 cowl at off-design condition.
• For CR 8.4 the peak value of pressure (~1.7x104 Pa) occurs at a location of 151mm, where as for CR 5.0 and 4.3 they occur at 188mm and 206mm having values ~1.6x104 Pa and ~1.4x104 Pa respectively. These locations indicate the likely locations of shock impingements inside the chamber.
• For cowl angle of 00 for a 10 mm chamber the maximum pressure value recorded is ~1.7x104 Pa whereas for 100 and 200 cowl it is ~1.1x104 Pa and 1.2 x104 Pa respectively. This is because in the first case the inlet is choked because of over contraction whereas in the other two cases the CR is less and flow is established inside the inlet.
• The average heat transfer rates of last four heat transfer gauges (180 mm, 190 mm, 200 mm and 210 mm from the forebody tip) for all lengths of cowls tested are found to be almost same (~ 20 W/cm2). This is because the flow is choked in all these cases. The numerical simulation also shows uniform distribution here, consistent with the experimental findings.
• The locations of heat transfer peaks for 100 cowl at design condition can be observed to be occurring at 170 mm and 200 mm from the forebody tip having values ~44 W/cm2 and ~39 W/cm2 respectively. For a 200 cowl they seem to be occurring at 170 mm and 180 mm from the forebody tip having values ~50 W/cm2 and ~30 W/cm2. These locations indicate the likely locations of shock impingements inside the chamber. With the evolution of concept of upstream fuel injection in recent times these may the most appropriate locations for fuel injection.
• At higher jet pressure ratios the plasma jet/ramp shock interaction results in a lambda shock pattern with the triple point forming vertically above the cowl level. This means the normal shock stands in front of the inlet making a part of the flow entering the inlet subsonic. The reflected shock from the triple point also separates the ramp boundary layer.
• At lower jet pressure ratios the triple point is formed below the cowl level and the flow entering inside the inlet is supersonic. The reflected shock interacts with the cowl shock and a weak separation shock is seen.
• Experiments are performed with concentrated DC electric discharge as energy source. Even though the amount of energy dumped here is less than 0.25% of the total energy it creates a perceptible disturbance in the flow.
• Experiments are also performed to see the effect of electric discharge as energy source on height of Mach stem for a given inlet configuration. Deposition of energy in the present location does not seem to alter the Mach stem height.
However more experiments need to be performed by varying the energy location to see its effect. Non-intrusive energy sources like microwave and lasers can be thought of as options for depositing energy to study its effect on Mach stem height. Since they provide more flexibility on varying the location of energy the optimum location of energy can be found out for highest reduction of Mach stem height.
47
(11355756), Jonathan J. Gaskins. "The Effect of Near Wall Disturbances on a Compressible Turbulent Boundary Layer." Thesis, 2021.
Abstract:
This study investigates the effects of near wall disturbances in the form of roughness on a compressible turbulent boundary layer. The studies were carried out using numerical methods which directly solve the Navier-Stokes equations. This provides for unique opportunities to investigate three dimensional structures of the flow as well as avoid the loss of physical fidelity with turbulence modeling. Three cases were ran, a smooth wall case, and two rough wall cases with different heights of the roughness elements between the cases. The results are first visualized with different approaches. Then statistical methods were used to characterize the flow.
48
Maity, Arnab. "Optimal Guidance Of Aerospace Vehicles Using Generalized MPSP With Advanced Control Of Supersonic Air-Breathing Engines." Thesis, 2012. https://etd.iisc.ac.in/handle/2005/2550.
Abstract:
A new suboptimal guidance law design approach for aerospace vehicles is proposed in this thesis, followed by an advanced control design for supersonic air-breathing engines. The guidance law is designed using the newly developed Generalized Model Predictive Static Programming (G-MPSP), which is based on the continuous time nonlinear optimal control framework. The key feature of this technique is one-time backward propagation of a small-dimensional weighting matrix dynamics, which is used to update the entire control history. This key feature, as well as the fact that it leads to a static optimization problem, lead to its computational efficiency. It has also been shown that the existing model predictive static programming (MPSP), which is based on the discrete time framework, is a special case of G-MPSP. The G-MPSP technique is further extended to incorporate ‘input inequality constraints’ in a limited sense using the penalty function philosophy. Next, this technique has been developed also further in a ‘flexible final time’ framework to converge rapidly to meet very stringent final conditions with limited number of iterations.
Using the G-MPSP technique in a flexible final time and input inequality constrained formulation, a suboptimal guidance law for a solid motor propelled carrier launch vehicle is successfully designed for a hypersonic mission. This guidance law assures very stringent final conditions at the injection point at the end of the guidance phase for successful beginning of the hypersonic vehicle operation. It also ensures that the angle of attack and structural load bounds are not violated throughout the trajectory. A second-order autopilot has been incorporated in the simulation studies to mimic the effect of the inner-loops on the guidance performance. Simulation studies with perturbations in the thrust-time behaviour, drag coefficient and mass demonstrate that the proposed guidance can meet the stringent requirements of the hypersonic mission.
The G-MPSP technique in a fixed final time and input inequality constrained formulation has also been used for optimal guidance of an aerospace vehicle propelled by supersonic air-breathing engine, where the resulting thrust can be manipulated by managing the fuel flow and nozzle area (which is not possible in solid motors). However, operation of supersonic air-breathing engines is quite complex as the thrust produced by the engine is a result of very complex nonlinear combustion dynamics inside the engine. Hence, to generate the desired thrust, accounting for a fairly detailed engine model, a dynamic inversion based nonlinear state feedback control design has been carried out. The objective of this controller is to ensure that the engine dynamically produces the thrust that tracks the commanded value of thrust generated from the guidance loop as closely as possible by regulating the fuel flow rate. Simultaneously, by manipulating throat area of the nozzle, it also manages the shock wave location in the intake for maximum pressure recovery with sufficient margin for robustness. To filter out the sensor and process noises and to estimate the states for making the control design operate based on output feedback, an extended Kalman filter (EKF) based state estimation design has also been carried out and the controller has been made to operate based on estimated states. Moreover, independent control designs have also been carried out for the actuators so that their response can be faster. In addition, this control design becomes more challenging to satisfy the imposed practical constraints like fuel-air ratio and peak combustion temperature limits. Simulation results clearly indicate that the proposed design is quite successful in assuring the desired performance of the air-breathing engine throughout the flight trajectory, i.e., both during the climb and cruise phases, while assuring adequate pressure margin for shock wave management.
49
Maity, Arnab. "Optimal Guidance Of Aerospace Vehicles Using Generalized MPSP With Advanced Control Of Supersonic Air-Breathing Engines." Thesis, 2012. http://etd.iisc.ernet.in/handle/2005/2550.
Abstract:
A new suboptimal guidance law design approach for aerospace vehicles is proposed in this thesis, followed by an advanced control design for supersonic air-breathing engines. The guidance law is designed using the newly developed Generalized Model Predictive Static Programming (G-MPSP), which is based on the continuous time nonlinear optimal control framework. The key feature of this technique is one-time backward propagation of a small-dimensional weighting matrix dynamics, which is used to update the entire control history. This key feature, as well as the fact that it leads to a static optimization problem, lead to its computational efficiency. It has also been shown that the existing model predictive static programming (MPSP), which is based on the discrete time framework, is a special case of G-MPSP. The G-MPSP technique is further extended to incorporate ‘input inequality constraints’ in a limited sense using the penalty function philosophy. Next, this technique has been developed also further in a ‘flexible final time’ framework to converge rapidly to meet very stringent final conditions with limited number of iterations.
Using the G-MPSP technique in a flexible final time and input inequality constrained formulation, a suboptimal guidance law for a solid motor propelled carrier launch vehicle is successfully designed for a hypersonic mission. This guidance law assures very stringent final conditions at the injection point at the end of the guidance phase for successful beginning of the hypersonic vehicle operation. It also ensures that the angle of attack and structural load bounds are not violated throughout the trajectory. A second-order autopilot has been incorporated in the simulation studies to mimic the effect of the inner-loops on the guidance performance. Simulation studies with perturbations in the thrust-time behaviour, drag coefficient and mass demonstrate that the proposed guidance can meet the stringent requirements of the hypersonic mission.
The G-MPSP technique in a fixed final time and input inequality constrained formulation has also been used for optimal guidance of an aerospace vehicle propelled by supersonic air-breathing engine, where the resulting thrust can be manipulated by managing the fuel flow and nozzle area (which is not possible in solid motors). However, operation of supersonic air-breathing engines is quite complex as the thrust produced by the engine is a result of very complex nonlinear combustion dynamics inside the engine. Hence, to generate the desired thrust, accounting for a fairly detailed engine model, a dynamic inversion based nonlinear state feedback control design has been carried out. The objective of this controller is to ensure that the engine dynamically produces the thrust that tracks the commanded value of thrust generated from the guidance loop as closely as possible by regulating the fuel flow rate. Simultaneously, by manipulating throat area of the nozzle, it also manages the shock wave location in the intake for maximum pressure recovery with sufficient margin for robustness. To filter out the sensor and process noises and to estimate the states for making the control design operate based on output feedback, an extended Kalman filter (EKF) based state estimation design has also been carried out and the controller has been made to operate based on estimated states. Moreover, independent control designs have also been carried out for the actuators so that their response can be faster. In addition, this control design becomes more challenging to satisfy the imposed practical constraints like fuel-air ratio and peak combustion temperature limits. Simulation results clearly indicate that the proposed design is quite successful in assuring the desired performance of the air-breathing engine throughout the flight trajectory, i.e., both during the climb and cruise phases, while assuring adequate pressure margin for shock wave management.
50
Kotnala, Sourabh. "Lattice Boltzmann Relaxation Scheme for Compressible Flows." Thesis, 2012. http://etd.iisc.ac.in/handle/2005/3257.
Abstract:
Lattice Boltzmann Method has been quite successful for incompressible
flows. Its extension for compressible (especially supersonic and hypersonic)
flows has attracted lot of attention in recent time. There have been some
successful attempts but nearly all of them have either resulted in complex
or expensive equilibrium function distributions or in extra energy levels.
Thus, an efficient Lattice Boltzmann Method for compressible fluid flows
is still a research idea worth pursuing for. In this thesis, a new Lattice
Boltzmann Method has been developed for compressible flows, by using the concept of a relaxation system, which is traditionally used as semilinear alternative for non-linear hypebolic systems in CFD. In the relaxation
system originally introduced by Jin and Xin (1995), the non-linear flux in a hyperbolic conservation law is replaced by a new variable, together with a relaxation equation for this new variable augmented by a
relaxation term in which it relaxes to the original nonlinear flux, in the limit of a vanishing relaxation parameter. The advantage is that instead of one non-linear hyperbolic equation, two linear hyperbolic equations need to be solved, together with a non-linear relaxation term. Based on the interpretation
of Natalini (1998) of a relaxation system as a discrete velocity Boltzmann equation, with a new isotropic relaxation system as the basic building block, a Lattice Boltzmann Method is introduced for solving the
equations of inviscid compressible flows. Since the associated equilibrium
distribution functions of the relaxation system are not based on a low Mach
number expansion, this method is not restricted to the incompressible limit.
Free slip boundary condition is introduced with this new relaxation system
based Lattice Boltzmann method framework. The same scheme is then extended
for curved boundaries using the ghost cell method. This new Lattice Boltzmann Relaxation Scheme is successfully tested on various bench-mark test cases for solving the equations of compressible flows such as shock tube problem in 1-D and in 2-D the test cases involving supersonic flow over a forward-facing step, supersonic oblique shock reflection from a flat plate, supersonic and hypersonic flows past half-cylinder.