Showing papers on "Supersonic speed published in 2008"
TL;DR: In this paper, a numerical study of mixing and combustion enhancement has been performed for a Mach 2 model scramjet (supersonic combustion ramjet) combustor Fuel (hydrogen) is injected at supersonic speed through the rear of a lobed strut located at the channel symmetry axis.
242 citations
TL;DR: In this paper, the existence of the bow shock was found to be dependent on the length of the nozzle's supersonic potential core, and the amount of standoff distance between the potential core and the substrate.
Abstract: Cold Spray involves the deposition of metallic powder particles using a supersonic gas jet. When the nozzle standoff distance is small, a bow shock is formed at the impingement zone between the supersonic jet and the substrate. It has long been thought that this bow shock is detrimental to process performance as it can reduce particle impact velocities. By using computational fluid dynamics, Particle Image Velocimetry and Schlieren imaging it was possible to show that the bow shock has a negative influence on deposition efficiency as a result of a reduction in particle velocity. Furthermore, the existence of the bow shock was shown to be dependent on the length of the nozzle's supersonic potential core. Experiments were carried out with aluminium, copper and titanium powders using a custom-made helium nozzle, operating at 2.0 MPa and 20 °C, and a commercial nitrogen nozzle operating at 3.0 MPa and 300 °C. In all cases, it was found that there is a direct relationship between standoff distance and deposition efficiency. At standoff distances less than 60 mm, the bow shock reduced deposition efficiencies by as much as 40%.
193 citations
TL;DR: The results of experimental and numerical investigations of the interaction between the near-wall electrical discharge and supersonic airflow in an aerodynamic channel with constant and variable cross sections are presented in this paper.
Abstract: The results of experimental and numerical investigations of the interaction between the near-wall electrical discharge and supersonic airflow in an aerodynamic channel with constant and variable cross sections are presented. Peculiar properties of the surface quasi-direct-current discharge generation in the flow are described. The mode with flow separation developing outside the discharge region is revealed as a specific feature of such a configuration. An interaction model is proposed on the basis of measurements and observations. A regime of gas-dynamic screening of a mechanical obstacle installed on the channel wall is studied. Variation of the main flow parameters caused by the surface discharge is quantified. The ability of the discharge to shift an oblique shock in an inlet is demonstrated experimentally. The influence of relaxation processes in nonequilibrium excited gas on the flow structure is analyzed. Comparison of the experimental data with the results of calculations based on the analytical model and on numerical simulations is presented.
139 citations
TL;DR: A reformulation of the refined similarity hypothesis in terms of the mass-weighted velocity rho1/3v yields scaling laws that are almost insensitive to the forcing of supersonic turbulent flow, implying that the most intermittent dissipative structures are shocks closely following the scaling of Burgers turbulence.
Abstract: The statistical properties of turbulence are considered to be universal at sufficiently small length scales, i.e., independent of boundary conditions and large-scale forces acting on the fluid. Analyzing data from numerical simulations of supersonic turbulent flow driven by external forcing, we demonstrate that this is not generally true for the two-point velocity statistics of compressible turbulence. However, a reformulation of the refined similarity hypothesis in terms of the mass-weighted velocity rho1/3v yields scaling laws that are almost insensitive to the forcing. The results imply that the most intermittent dissipative structures are shocks closely following the scaling of Burgers turbulence.
125 citations
TL;DR: In this article, a weakly non-parallel flow analysis was proposed to eliminate the critical layer singularity of the SINR model, which has a strong effect on the shape of the peak noise spectrum.
Abstract: This paper is concerned with utilizing the acoustic analogy approach to predict the sound from unheated supersonic jets. Previous attempts have been unsuccessful at making such predictions over the Mach number range of practical interest. The present paper, therefore, focuses on implementing the necessary refinements needed to accomplish this objective. The important effects influencing peak supersonic noise turn out to be source convection, mean flow refraction, mean flow amplification, and source non-compactness. It appears that the last two effects have not been adequately dealt with in the literature. The first of these because the usual parallel flow models produce most of the amplification in the so called critical layer where the solution becomes singular and, therefore, causes the predicted sound field to become infinite as well. We deal with this by introducing a new weakly non parallel flow analysis that eliminates the critical layer singularity. This has a strong effect on the shape of the peak noise spectrum. The last effect places severe demands on the source models at the higher Mach numbers because the retarded time variations significantly increase the sensitivity of the radiated sound to the source structure in this case. A highly refined (non-separable) source model is, therefore, introduced in this paper.
110 citations
TL;DR: In this article, the authors consider the problem of two-dimensional supersonic flow onto a solid wedge, or equivalently in a concave corner formed by two solid walls, and show that the timedependent solution is self-similar, with a weak shock at the tip of the wedge.
Abstract: We consider the problem of two-dimensional supersonicflow onto a solid wedge, or equivalently in a concave corner formed by two solid walls. For mild corners, there are two possible steady state solutions, one with a strong and one with a weak shock emanating from the corner. The weak shock is observed in supersonic flights. A longstanding natural conjecture is that the strong shock is unstable in some sense. We resolve this issue by showing that a sharp wedge will eventually produce weak shocks at the tip when accelerated to a supersonic speed. More precisely, we prove that for upstream state as initial data in the entire domain, the timedependent solution is self-similar, with a weak shock at the tip of the wedge. We construct analytic solutions for self-similar potential flow, both isothermal and isentropic with arbitrary� � 1. In the process of constructing the self-similar solution, we develop a large number of theoretical tools for these elliptic regions. These tools allow us to establish large-data results rather than a small perturbation. We show that the wave pattern persists as long as the weak shock is supersonic-supersonic; when this is no longer true, numerics show a physical change of behavior. In addition, we obtain rather detailed information about the elliptic region, including analyticity as well as bounds for velocity components and shock tangents. c � 2007 Wiley Periodicals, Inc.
97 citations
TL;DR: In this paper, a simulation of a rectangular, mixed-compression inlet has been performed on a 20 x 10 6 points mesh using the delayed detached-eddy simulation method, a version of detachededdy simulation that ensures the attached boundary layers are treated using Reynolds-averaged Navier-Stokes equations.
Abstract: Supersonic inlet buzz in a rectangular, mixed-compression inlet has been simulated on a 20 x 10 6 points mesh using the delayed detached-eddy simulation method, a version of detached-eddy simulation that ensures the attached boundary layers are treated using Reynolds-averaged Navier-Stokes equations. The results are compared with experimental data obtained during a previous campaign of wind-tunnel experiments. The comparison of unsteady data is performed thanks to phase averages, Fourier transforms, and wavelet transforms. The buzz observed at Mach 1.8, which occurred at a frequency of 18 Hz, is well reproduced. The shock oscillations, as well as the different flow features experimentally observed, are present in the simulation. The buzz frequency, as well as higher frequencies existing in the experimental pressure signals, are correctly predicted. The data issued from the simulation (time history of pressure fluctuations, pseudo-Schlieren, and three-dimensional visualizations) allow a better investigation of the inlet flowfield during buzz and a detailed description and physical analysis of this phenomenon. A description and an explanation of the mechanism at the origin of secondary oscillations that occur at a higher frequency during buzz are proposed. The crucial role of acoustic waves moving through the duct is shown.
96 citations
TL;DR: In this paper, the authors review the physics basis of our present understanding of the transition process and address the need for careful experimental work as per the guidelines enunciated years ago by the U.S. Transition Study Group.
Abstract: *Much has been learned about the physics underlying the transition process at supersonic and hypersonic speeds through years of analysis, experiment and computation. The application has been principally to simple shapes like plates, cones and spherical nosetips. But the shapes of the new entry vehicles are not simple. They will invariably be at angle of attack so three dimensional effects will be very important, as will roughness effects due to ablation. This paper will review the physics basis of our present understanding of the transition process. Further, because of the complex geometries, it will address the need for careful experimental work as per the guidelines enunciated years ago by the U.S. Transition Study Group. Following these guidelines is essential to obtaining reliable, usable data for use in refining transition estimation techniques . I. Introduction Entry vehicles descend rapidly through the atmosphere and decelerate as the drag forces increase with the increasing atmospheric density. The vehicle starts out fully laminar. Since Reynolds numbers increase rapidly through the des cent, transition tends to move very rapidly over the whole vehicle over a narrow range of altitudes. O ne tries to minimize the aerodynamic heating loads in entry so as to minimize the thermal protection needs of the vehicle. This means delaying transition to as low an altitude as possible . The history of high -performance entry vehicles begins with the development of the ICBM in the 1950’s. These v ehicles were essentially cone -cylinders with large nose bluntness to minimize stagnation point heating. The nose materials were often subliming ablators to take advantage of the further reduction in stagnation region heating due to the surface blowing from the subliming surface. Transition would move forward from the cylinder to the cone as the vehicle descende d through the atmosphere. If transition occurred asymmetrically on the cylinder or cone (leading to drag asymmetry), it was essential that the asymmetric effects be small enough to be corrected by the control system of the missile so as to minimize the cir cle of error about the target. In some cases, the transition unexpectedly moved forward onto the spherical nose , a phenomenon named the “blunt -body -paradox.” A study in that time period by Allen and Eggers 1 showed that for orbital and sub -orbital entry ( < 8 km/sec) , the convective aerodynamic heating rate was directly related to the vehicle’s ballistic coefficient, (W/C DA). The lower the ballistic coefficient, the lower the heating rates in entry. Hence the tendency to low weight and high drag coefficient. The highly blunted capsul e shapes of the Mercury, Gemini, Apollo and Soyuz vehicles with a blat ing heat shields are a consequence of this argument . At entry speeds above about 10 km/sec , the shock layers ahead of the entering blunt shapes become lumino us and radiate. These radiative heat transfer rates are highly density dependent. Thus the blunt body may not be the best entry shape for supercircular entry speeds. Studies by Allen et al 2,3 show that to minimize total heat transfer or total ablated mass , the optimum shape gets progressively slenderer as the entry speed is increased. Not enough was done with these vehicles to identify the major transition issues.
95 citations
TL;DR: In this article, the authors examined supersonic microjets of 100-1,000 μm in size with exit velocities in the range of 300-500 m/s.
Abstract: The fluid dynamics of microflows has recently commanded considerable attention because of their potential applications. Until now, with a few exceptions, most of the studies have been limited to low speed flows. This experimental study examines supersonic microjets of 100–1,000 μm in size with exit velocities in the range of 300–500 m/s. Such microjets are presently being used to actively control larger supersonic impinging jets, which occur in STOVL (short takeoff and vertical landing) aircraft, cavity flows, and flow separation. Flow properties of free as well as impinging supersonic microjets have been experimentally investigated over a range of geometric and flow parameters. The flowfield is visualized using a micro-schlieren system with a high magnification. These schlieren images clearly show the characteristic shock cell structure typically observed in larger supersonic jets. Quantitative measurements of the jet decay and spreading rates as well as shock cell spacing are obtained using micro-pitot probe surveys. In general, the mean flow features of free microjets are similar to larger supersonic jets operating at higher Reynolds numbers. However, some differences are also observed, most likely due to pronounced viscous effects associated with jets at these small scales. Limited studies of impinging microjets were also conducted. They reveal that, similar to the behavior of free microjets, the flow structure of impinging microjets strongly resembles that of larger supersonic impinging jets.
93 citations
TL;DR: In this paper, the behavior of high-pressure natural gas in supersonic nozzles is studied and compared with the perfect gas case. But the results show a significant variation in gas properties estimation.
Abstract: The computational fluid dynamics technique was used to study the behavior of high-pressure natural gas in supersonic nozzles. Although many applications of gas flow produce insignificant errors when the gas is assumed ideal, our results indicate significant variation of gas properties. This article illustrates natural gas behavior when it is considered to be real and how erroneous the properties may become when the gas is assumed to be ideal. The article also presents the influences of properties related to the flow of natural gas through supersonic nozzles. Using a quite accurate equation of state model, real gas effects are studied and compared with the perfect gas case. The results show a significant variation in gas properties estimation. Location of shockwave is also analyzed. The comparison of results for two gases (methane and nitrogen) indicated that shockwave position can significantly change when the gas is considered as real rather than perfect.
86 citations
TL;DR: In this article, the importance of discrete roughness and the correlations developed to predict the onset of boundary layer transition on hypersonic flight vehicles are discussed and compared to the ground-based correlations.
Abstract: The importance of discrete roughness and the correlations developed to predict the onset of boundary layer transition on hypersonic flight vehicles are discussed. The paper is organized by hypersonic vehicle applications characterized in a general sense by the boundary layer: slender with hypersonic conditions at the edge of the boundary layer, moderately blunt with supersonic, and blunt with subsonic. This paper is intended to be a review of recent discrete roughness transition work completed at NASA Langley Research Center in support of agency flight test programs. First, a review is provided of discrete roughness wind tunnel data and the resulting correlations that were developed. Then, results obtained from flight vehicles, in particular the recently flown Hyper-X and Shuttle missions, are discussed and compared to the ground-based correlations.
TL;DR: In this paper, the effects of some of these parameters that appear to determine control efficiency are examined and some of the fundamental mechanisms behind this control approach are explored and it has been clearly demonstrated that the activation of microjets leads to a local thickening of the jet shear layer, near the nozzle exit, making it more stable and less receptive to disturbances.
Abstract: Supersonic impinging jet(s) inherently produce a highly unsteady flow field. The occurrence of such flows leads to many adverse effects for short take-off and vertical landing (STOVL) aircraft such as: a significant increase in the noise level, very high unsteady loads on nearby structures and an appreciable loss in lift during hover. In prior studies, we have demonstrated that arrays of microjets, appropriately placed near the nozzle exit, effectively disrupt the feedback loop inherent in impinging jet flows. In these studies, the effectiveness of the control was found to be strongly dependent on a number of geometric and flow parameters, such as the impingement plane distance, microjet orientation and jet operating conditions. In this paper, the effects of some of these parameters that appear to determine control efficiency are examined and some of the fundamental mechanisms behind this control approach are explored. Through comprehensive two- and three-component velocity (and vorticity) field measurements it has been clearly demonstrated that the activation of microjets leads to a local thickening of the jet shear layer, near the nozzle exit, making it more stable and less receptive to disturbances. Furthermore, microjets generate strong streamwise vorticity in the form of well-organized, counter-rotating vortex pairs. This increase in streamwise vorticity is concomitant with a reduction in the azimuthal vorticity of the primary jet. Based on these results and a simplified analysis of vorticity transport, it is suggested that the generation of these streamwise vortices is mainly a result of the redirection of the azimuthal vorticity by vorticity tilting and stretching mechanisms. The emergence of these longitudinal structures weakens the large-scale axisymmetric structures in the jet shear layer while introducing substantial three-dimensionality into the flow. Together, these factors lead to the attenuation of the feedback loop and a significant reduction of flow unsteadiness.
TL;DR: In this article, the authors used the computational fluid dynamics technique to study the behavior of high-pressure natural gas when it flows through nozzles with supersonic velocities.
Abstract: The computational fluid dynamics technique is used to study the behavior of high-pressure natural gas when it flows through nozzles with supersonic velocities. Effect of nozzle geometry is discussed by inserting a constant area channel between the convergent and divergent parts of the system. Various conduit lengths are analyzed to show how the minimum temperature could be influenced by the geometry of the nozzle. The results also show that changing channel length can affect the position of shockwave. The results for the effect of vorticity on the performance of the nozzles show that, although losses in pressure increase due to inlet swirl flow, vorticity increases very sharply in the vicinity of the shock. It could be concluded that the region just before the shock spot is the main region where fine particles can be separated because of the large vorticity strength. Shock with reasonable strength may be favored in practical applications where fine particles separation is desired.
TL;DR: An overview of key findings obtained from computational studies of supersonic micronozzle flow are provided and the implications for future micro-scale nozzle design and optimisation are discussed.
Abstract: The next-generation of small satellites ('nanosats') will feature masses <10 kg and require miniaturised propulsion systems capable of providing extremely low levels of thrust The emergence of viscous, thermal and/or rarefaction effects on the micro-scale can significantly impact the flow behaviour in supersonic micronozzles resulting in performance characteristics which differ substantially from traditional macro-scale nozzle designs In this paper, we provide an overview of key findings obtained from computational studies of supersonic micronozzle flow and discuss the implications for future micro-scale nozzle design and optimisation
TL;DR: In this article, an experimental study showed the possibility of activating plasma discharges at the tip of a supersonic projectile flying in conditions encountered at a low altitude using highvoltage generators.
Abstract: Experimental and theoretical investigations on the possibilities of steering a supersonic projectile by using a plasma actuator started in 2001, but they have not been published up to now, for confidentiality reasons. The experimental study shows the possibility of activating plasma discharges at the tip of a supersonic projectile flying in conditions encountered at a low altitude. Plasma discharges were produced by the use of high-voltage generators that were able to supply electric discharges between two electrodes flush with the conical surface of the projectile nose. Visualizations show that the generation of a plasma discharge produces a perturbation between the projectile surface and the shock wave attached to the conical projectile tip. The perturbation is strong enough to distort the shock wave. A numerical simulation was performed for an ideal gas, in which the plasma discharge was modeled as a transverse hot jet. The comparison between the flow visualizations and the numerical results shows the similarity between the visualized and the computed flow structures. The results show that the asymmetry of the flowfield around the projectile produces a lateral force and a pitching moment that favorably combine to steer the projectile.
TL;DR: In this article, the authors studied the transonic shock problem for the Euler flows through a class of 2-D or 3-D nozzles, where the nozzle is assumed to be symmetric in the diverging (or converging) part.
TL;DR: In this paper, calculations of flow within the two-dimensional Euler model of supersonic swirling flow of gas in a super-separator of natural gas are given, numerical experiment is performed, and the basic parameters of gas flow (velocity components, pressure, and so on) are obtained as functions of radius.
Abstract: Results are given of calculations of flow within the two-dimensional Euler model of supersonic swirling flow of gas in a supersonic separator of natural gas. The formulation of the problem is given, numerical experiment is performed, and the basic parameters of gas flow (velocity components, pressure, and so on) are obtained as functions of radius. The process of relaxation of flow to steady state with the formation of shock wave is considered, and the shock wave structure is determined. The behavior of gasdynamic parameters is analyzed under conditions of separation in the region of shock wave and behind it.
21 Jul 2008
TL;DR: In this article, a new scramjet engine model has been developed to support hypersonic vehicle design studies and flight dynamics and control system analysis, which is suitable for use with a control-oriented dynamic model of a hypercarrier.
Abstract: A new scramjet engine model has been developed to support hypersonic vehicle design studies and flight dynamics and control system analysis. This paper explains the methodology and the governing equations for the new propulsion system model that is suitable for use with a control oriented dynamic model of a hypersonic vehicle. Previous propulsion models used for this purpose were based on simple Rayleigh flow for the combustion process, but despite this, captured the propulsion system interactions with the vehicle aerodynamics and structural dynamics. A new, higher fidelity propulsion system model is constructed that simulates numerous phenomena that were neglected in the Rayleigh flow approach. The new model is of higher fidelity, and therefore it is not designed to calculate the flow physics on a timescale that is suitable for dynamics and control simulations. Instead it will be used as a truth model and the starting point for the derivation of a reduced-order model. Specific phenomena that are included in the new model are: a pre-combustion shock train within the isolator and its interactions with the combustor, the loss of stagnation pressure due to gas dissociation and recombination, wall heat transfer and skin friction, a fuel-air mixing submodel, and a finite-rate chemistry and autoignition reaction mechanism. It is shown that the new propulsion system model expands the operability envelope as compared to the previous model by accommodating ramjet combustion, which occurs at high supersonic/low hypersonic flight Mach numbers.
TL;DR: In this article, a self-ignited supersonic combustion experiment was performed using a cavity-based injector in the T3 free-piston shock tunnel, using various combustor inlet and fuel-flow conditions.
Abstract: Self-ignited supersonic combustion experiments have been performed using a cavity-based injector in the T3 free-piston shock tunnel, using various combustor inlet and fuel-flow conditions. Planar laser-induced fluorescence on the hydroxyl radical and fast-acting pressure transducers are used to investigate the flow characteristics. Four hydrogen injectors are located upstream of an open cavity. The separated shear layer reattaches, generating an oblique shock at the cavity's trailing edge and establishing the major flow structure. The normalized pressure rise due to combustion increases as the equivalence ratio increases and the freestream stagnation enthalpy decreases, over the range of conditions' tested. Angled injection upstream of the cavity allows the cavity to act as a flame holder. High injection pressure helps to ignite immediately upstream of the injector and forms two flame layers over the cavity. The fluorescence peak signal shows periodic maxima near the cavity, and the interval between peaks decreases as the equivalence ratio is increased. Low-total-enthalpy conditions also exhibit longer ignition-delay distances. Comparison of fluorescence images and static pressure measurements indicates that, at these conditions, the heat release is mostly initiated by the shock wave from the cavity's trailing face and the ignition above the cavity does not have a strong influence on the downstream combustion.
TL;DR: In this paper, the development of isolated turbulent spots in supersonic flat plate boundary layers by direct numerical simulations is investigated and mean velocity profiles and Reynolds stresses are determined by averaging, for one flow parameter set, over an ensemble of simulations.
TL;DR: In this article, a simulation of a two-dimensional supersonic linear micronozzle with expansion angles larger than traditional macro-scale nozzles has been performed, and it was found that an inherent tradeoff exists between the viscous losses and the losses resulting from the nonaxial exit flow at larger expansion angles.
Abstract: A comprehensive numerical investigation of a steady viscous flow through a two-dimensional supersonic linear micronozzle has been performed. The baseline model for the study is derived from the NASAGoddard Space Flight Center microelectromechanical systems-based hydrogen peroxide prototype microthruster. On the microscale, substantial viscous subsonic layers may form on the nozzle expander walls, which reduce thrust and efficiency. One approach to compensate for the presence of these layers has been to designmicronozzleswith expansion angles larger than traditional macroscale nozzles. Numerical simulations have been conducted for a range of Reynolds numbers (Re 15–800) and for expander half-angles of 10–50 deg. Twodifferentmonopropellant fuels have been considered: decomposed 85% pure hydrogen peroxide and decomposed hydrazine. It is found that an inherent tradeoff exists between the viscous losses and the losses resulting from the nonaxial exit flow at larger expansion angles. Our simulations indicate that themaximumnozzle efficiency occurs for bothmonopropellants at a nozzle expansion halfangle of approximately 30 deg, which is significantly larger than that of traditional conical nozzle designs.
TL;DR: In this paper, a short-pulse repetitive discharge is used to ignite hydrogen jet flames in supersonic crossflows, where a nonequilibrium plasma is produced by repetitive pulses of 7-kV peak voltage, 20-ns pulsewidth, and 50-kHz repetition rate.
Abstract: A short-pulse repetitive discharge is used to ignite hydrogen jet flames in supersonic crossflows. Nonequilibrium plasma is produced by repetitive pulses of 7-kV peak voltage, 20-ns pulsewidth, and 50-kHz repetition rate. Sonic or subsonic hydrogen jets are injected into a pure-oxygen supersonic free-stream flow of Mach numbers M = 1.7-2.3. The fuel injection nozzles and electrodes are mounted flush with the surface of a flat plate that is oriented to be parallel to the flow to minimize stagnation pressure losses associated with generated shock waves. A configuration combining an upstream subsonic oblique jet and a downstream sonic transverse jet serves to provide an adequate flow condition for jet flame ignition. The flow pattern and shock waves induced by the dual hydrogen jets are characterized by Schlieren imaging. Planar-laser-induced fluorescence and emission spectroscopy are employed for imaging the distribution of OH radicals. The OH fluorescence image of the region in the vicinity of the discharge confirms jet flame ignition by the plasma.
21 Jul 2008
TL;DR: In this article, high-order compact differencing/filtering schemes are coupled with recently developed localized artificial diffusivity methodology in the context of large-eddy simulation (LES) to obtain insights into the physics of an under-expanded sonic jet injection into a supersonic crossflow.
Abstract: High-order compact differencing/filtering schemes are coupled with recently developed localized artificial diffusivity methodology in the context of large-eddy simulation (LES) to obtain insights into the physics of an under-expanded sonic jet injection into a supersonic crossflow. The flow conditions of the experiment by Santiago and Dutton [J. Prop. Power. 13 (1997) 264–273] are selected for detailed simulation. The present LES qualitatively reproduce the unsteady dynamics of both barrel shock and bow shock as observed in the experiment. It found that pressure fluctuation inside the upstream recirculation region induces unsteadiness of windward jet shear layer and causes large-scale dynamics of the barrel shock and front bow shock. Statistics obtained by the LES also show good agreement with the experiment. With regard to the processes controlling the jet mixing we studied the dynamics of vortex structures in the flow. Two regions of vortex formation which form hairpin-like structure are identified in the windward and leeward jet boundaries. These vortices play an important role in determining the behavior of jet fluid stirring and subsequent mixing. Also noted is the entrainment of rolled-up windward shear layer by the upstream recirculating flow.
TL;DR: In this article, a short selective review of theoretical and experimental studies conducted by the authors and their collaborators during the past few years in areas related to supersonic and hypersonic flow regimes with applications such as drug reduction, inlet and effective vehicle geometry control in off-design flight regimes, and steering and sonic boom mitigation is presented.
Abstract: This paper presents a short selective review of theoretical and experimental studies conducted by the authors and their collaborators during the past few years in areas related to supersonic and hypersonic flow regimes with applications such as drug reduction, inlet and effective vehicle geometry control in off-design flight regimes, and steering and sonic boom mitigation. Their results suggest a principal possibility to enable transitions between the propulsion modes and ramjet startup and to minimize the need for the traditional isolator stage, as well as to increase the inlet mass capture at Mach numbers below the design value, using active control based on virtual shapes created by energy addition upstream of the inlet throat. A common feature of all these substantially different applications and processes is a power deposition into a supersonic flow which results in the creating of virtual shapes, modifying flowlike solid obstacles immersed in it. The virtual shapes can be created by microwave plasma heating, magnetohydrodynamic forces, electron beams, and localized plasma-assisted surface combustion. The power necessary to operate plasmas can come either from the turbine at mode transition, from an auxiliary power unit, or, as suggested in a new bypass concept, from a magnetohydrodynamic generator either placed downstream of the combustor or collocated with it.
TL;DR: In this paper, a full-scale scramjet combustor with kerosene fuel injected from struts placed in the combustor flowpath was designed and analyzed using computational-fluid-dynamics software.
Abstract: Computational-fluid-dynamics-based design and analysis is presented for a full-scale scramjet combustor with kerosene fuel injected from struts placed in the combustor flowpath. Three-dimensional Navier-Stokes equations are solved with a K-e turbulence model using commercial computational-fluid-dynamics software. Combustion is modeled based on infinitely fast chemical kinetics. Lagrangian dispersed-phase analysis is considered for fueldroplet evaporation and mixing in the supersonic stream. Parametric studies are carried out to investigate the effect of combustor-inlet Mach number and total pressure on the flow development process. A higher combustor-entry Mach number and distributed-fuel-injection system will ensure the existence of predominant supersonic flow in the combustor. Simulations are also carried out to investigate two different kinds of fuel injection struts in the scramjet combustor performance. A distributed-fuel -injection system, required to avoid thermal choking, increases the three-dimensionality of the flowfield.
TL;DR: In this paper, the effects of the termination shock in neutrino-driven winds and its role in the r-process were investigated, and it was shown that the slowdown of the temperature decrease plays a decisive role in determining the R-process abundance curves.
Abstract: Recent hydrodynamic studies of core-collapse supernovae imply that the neutrino-heated ejecta from a nascent neutron star develop to supersonic outflows. These supersonic winds are influenced by the reverse shock from the preceding supernova ejecta, forming the wind termination shock. We investigate the effects of the termination shock in neutrino-driven winds and its role in the r-process. Supersonic outflows are calculated with a semianalytic neutrino-driven wind model. Subsequent termination-shocked, subsonic outflows are obtained by applying the Rankine-Hugoniot relations. We find a couple of effects that can be relevant for the r-process. First is the sudden slowdown of the temperature decrease by the wind termination. Second is the entropy jump by termination-shock heating, up to several hundred NAk. Calculations of nucleosynthesis in the obtained winds are performed to examine these effects on the r-process. We find that the slowdown of the temperature decrease plays a decisive role in determining the r-process abundance curves. This is due to the strong dependences of the nucleosynthetic path on the temperature during the r-process freezeout phase. Our results suggest that only the termination-shocked winds with relatively small shock radii (~500 km) are relevant for the bulk of the solar r-process abundances (A ≈ 100-180). The heaviest part of the solar r-process curve (A ≈ 180-200), however, can be reproduced in both shocked and unshocked winds. These results may help to constrain the mass range of supernova progenitors relevant for the r-process. We find, on the other hand, a negligible role of the entropy jump in the r-process. This is because the sizable entropy increase takes place only at a large shock radius (10,000 km), where the r-process has already ceased.
TL;DR: In this article, the energy loss of a uniformly moving dislocation radiating lattice waves is quantitatively studied by atomistic simulations and the velocity dependence of the energy radiation provides a unique way to understand limiting behavior exhibited by fast moving dislocations.
Abstract: The energy loss of a uniformly moving dislocation radiating lattice waves is quantitatively studied by atomistic simulations. The velocity dependence of the energy radiation provides a unique way to understand limiting behavior exhibited by fast moving dislocations. In combination with theoretical analyses, the simulations provide a consistent atomic-level picture of near-sonic, transonic, and supersonic dislocation motions in crystalline materials.
07 Jan 2008
TL;DR: The Gulfstream Quiet Spike as discussed by the authors is a telescoping forward fuselage extension that alters the bow shock of the classic N-wave pressure signature generated by aircraft traveling at supersonic speeds.
Abstract: The Gulfstream Quiet Spike is a telescoping forward fuselage extension that alters the bow shock of the classic N-wave pressure signature generated by aircraft traveling at supersonic speeds. The bow shock is broken up into sequence of weak shocks that propagate parallel to each other without coalescing. The resulting rise time to peak overpressure in the ground signature is increased from two or three milliseconds to thirty or forty milliseconds, significantly reducing the loudness. The development history of the Quiet Spike concept is traced from initial conception up to the design of a flight test configuration. Early aerodynamic and structural tests matured the concept from one viewed with skepticism to one deemed possible. These tests and the subsequent development of an aerodynamic configuration that could be installed on an F-15B test aircraft are summarized and the objectives of the flight test program are discussed.
TL;DR: In this paper, the fractal dimension of the transitional region and the fully developing turbulence region of the supersonic mixing layer were measured based on the box-counting method, and the corresponding images distinctly reproduced the flow structure of laminar, transitional and turbulent region.
Abstract: Flow visualization of supersonic mixing layer has been studied based on the high spatiotemporal resolution Nano-based Planar Laser Scattering (NPLS) method in SML-1 wind tunnel. The corresponding images distinctly reproduced the flow structure of laminar, transitional and turbulent region, with which the fractal measurement can be implemented. Two methods of measuring fractal dimension were introduced and compared. The fractal dimension of the transitional region and the fully developing turbulence region of supersonic mixing layer were measured based on the box-counting method. In the transitional region, the fractal dimension will increase with turbulent intensity. In the fully developing turbulent region, the fractal dimension will not vary apparently for different flow structures, which embodies the self-similarity of supersonic turbulence.
TL;DR: In this article, a twisted conical wire array is used to produce convergent plasma flows each rotating about the central axis, and collision of the flows produces a standing shock and jet that each have supersonic azimuthal velocities.
Abstract: The first laboratory astrophysics experiments to produce a radiatively cooled plasma jet with dynamically significant angular momentum are discussed. A new configuration of wire array $z$ pinch, the twisted conical wire array, is used to produce convergent plasma flows each rotating about the central axis. Collision of the flows produces a standing shock and jet that each have supersonic azimuthal velocities. By varying the twist angle of the array, the rotation velocity of the system can be controlled, with jet rotation velocities reaching $\ensuremath{\sim}18%$ of the propagation velocity.