28th International Symposium on Shock Waves

The paper is devoted to numerical simulation of processes accompanying underwater explosion near the free surface. In spite of the fact that the problem has been attracting attention of many researchers (see, for example, [1, 2, 3]) challenging problems still remain. The main attention in this paper is paid to compressible and rarefaction wave propagation, to their interaction with each other and with the liquid-gas interface, to deformation of the free surface.

Analysis of these phenomena is important for both progress of theory of heterogeneous media and applications utilizing specific features of underwater explosions. One of goals of this study is development of efficient, convenient, and flexible tool for investigations of such phenomena.

Shock wave interaction with granular media has attracted great interest due to its relevance to many practical applications such as shock attenuation by granular filters [1] and powder densification and consolidation by shock waves [2]. The primary concern of recent investigations on this phenomenon is the stress enhancement by a granular material placed on a solid wall, instead of expected stress damping. Suzuki and Adachi [3] observed that the propagation of a shock wave over a solid wall covered with a thin dust layer gave rise to a peak in pressure profiles measured under the dust layer. Gelfand

et al

. [4] also verified that when a granular layer covered the end-wall of a vertical shock tube, the pressure measured at the end-wall after headon collision of a shock wave with the granular layer temporally became larger than that without granular material. In order to understand the fundamental mechanism of the stress enhancement, numerous experimental works ([5]–[12]) have been carried out using vertical shock tubes similar to that used by Gelfand

et al

.

The direct applications of water exit and water entry study are submarine launched ballistic missile and anti-submarine missile [1],[2]. This study is also related to underwater high-speed torpedo which moves in a supercavity and is designed to confront aircraft carrier [3]-[4]. Previous research has suggested that when an underwater body moves close to the free surface, cavitation may become important[5],[6]. However, due to its transient and non-linear nature,water exit is a rather complicated process and many problems remain to be solved.

Obviously, it is beyond question that the explosive character of decompression during the volcanic eruption initiates a whole spectrum of phenomena in a precompressed magmatic melt containing large amounts of dissolved gases: homogeneous nucleation, bubbly cavitation, gas diffusion, and dynamically increasing viscosity of the melt. It is these processes that determine the eruption character, the magma state dynamics, and the flow structure in decompression waves as a whole. In the same time many aspects of their mechanisms remain unclear. The answers to these questions and, in particular, to the question about the mechanism of the cavitating magma transition to a state of an ash cloud cannot be simple because of extremely complicated and multiple-scale phenomena such as an explosive eruption [1]. In this connection, as it was noted by Gilbert and Sparks [2], laboratory experiments on the dynamics researches of volcanic flows and, in particular, the shock tube methods must become important components of simulation processes together with mathematical models and numerical analysis.

The rapid release of energy beneath the surface of water results in movement of the surrounding water, especially on the side of free surface. This can be realized using chemical and nuclear explosions, electric discharge, or pulsed laser focusing. By carefully setting the depth of explosion in a confined tube, a pulsed water jet can be effectively generated at the exit of the tube. The laser-induced liquid jet has been studied in biology and medicine [1]. This water jet is also important in volcanology to investigate explosive volcano eruptions that are driven by expanding gas bubbles [2]. In these applications, the bubble is generated in a confined space as in a pipe or a channel. Shallow and deep underwater explosions have been heavily investigated experimentally and numerically for decades. However the behavior of collapsing bubble and water jet in a confined space have not much been studied so far. The motivation of this study is to investigate the behavior of collapsing bubble and water jet formation in a narrow tube with a rectangular cross-section by underwater explosion. Special attentions are paid on the repeatability, so that the collected data can be used to validate numerical models for these compressibility dominated two-phase flows.

When a shock wave propagates over many small solid particles on a horizontal wall, some particles near the surface of the layer are lifted and dispersed into the shockinduced flow. These dispersed particles is called the dust cloud. This phenomenon is actually seen in galleries of coal mines or in pipelines for neumatic transportation of powder, and mixing dispersed flammable dust particles with high-temprature and high-pressure gas behind the shock wave sometimes causes the dust explosion. And this phenomenon includes some interesting factors, such as the shock structures interacted with the dust layer, interactions between gas and solid particle, and interactions between solid particles.

The pressure histories obtained when a shock wave propagates into an air-solid particle medium is well known: the overpressure jump decreases, as the shock wave propagates into the mixture and is followed by a pressure build-up corresponding to the velocity relaxation processes. In the present paper, an air-water droplet mixture interacting with a shock wave has been studied and the comportment of the pressure traces was found significantly changed in comparison to the interaction with a air-solid particle mixture. This is attributed to the ability of the droplets to deform and fragment into finer ones. This phenomenon, known as secondary atomisation, widely reviewed by Gelfand[1] and by Guildenbecher[2], affects both the pressure histories and the impulse induced by the shock wave. We have previously studied the influence of the height of cloud of droplets on shock wave propagation [3]. In the present work, we focus our attention on the influence of the droplet diameter on the attenuation of shock wave propagating into the air-water mixture. Moreover, predictions obtained by 1D numerical simulations are compared to the experimental results. The necessity to introduce a secondary atomisation model to fit the experimental behaviour is then underlined.

A hypersonic vehicle travelling through weather may encounter a wide range of particle environments. The impact of these particles on the vehicle surface may pose a significant erosion hazard to the vehicle’s thermal protection system. However, interaction with the flowfield will alter the particle motion, mitigating the erosive impact.

Multimaterial problems have always been a challenging topic for research, due to both their complexity, and the range of applications concerned. Several ways of modelling have been developed for such problems during the last decades. From shock impacts between compressible fluids to fluid-structure interaction, with elasto-plastic deformations, several problems may arise. The relevant methods have to take into account the different behaviours that the materials can exhibit, whilst conserving a sharp and accurate interface. In this paper, a new 3D conservative method for interface tracking based on level-set functions is adressed. Each material is treated independently except at the interface where appropriate boundary conditions need to be specified.Solving the Riemann problem at the interface depends on the different set of equations for each component. A new Riemann solver dedicated on solid/fluid interactions has also been derived. Fluid behaviours are governed by the Euler equations, while a recently developed non-linear elasto-plastic model will be used for solids. The sharp interface method is based upon a strict finite volume evaluation of the governing constitutive laws on fixed Cartesian meshes.

Shock waves in two-phase compressible flows are a fundamental topic in science and engineering. To better understand instability phenomena that are important for the evolution of such flows, basic configurations such as shock-bubble interactions in two-phase compressible flows are considered to investigate the Richtmyer-Meshkov instability and Rayleigh-Taylor instability. Flows of this type are present in many engineering applications including supersonic mixing and combustion systems and extra-corporeal shock-wave lithotripsy.

Gas-particle two-phase flows are abundant in the practical engineering problems. A branch of application is involved with the solid particles such as dust cyclone, flight in the hail storm, detonation of the dusty gases in the coal mine and grain storage tank, etc. Another branch of applications happen in the gas-liquid droplet two-phase flows but it is somewhat strange that they have not been much solved in the literature. Examples of this flow are the aviation of a flight vehicle through the cloud, fuel spray combustion and pulse detonation engine. Numerical simulation of the gas flow with the liquid droplets is more difficult than the gas flow with the solid particles. The reason is simple: the liquid droplets are a deformable medium with surface tension, breakup, coalescence, evaporation and condensation. In the present paper, we numerically simulate the transonic flow of air-water droplet mixture over a NACA0012 airfoil: see Fig. 1.

The collapse of cavitation bubbles near walls is one of the major reasons for failure of technical devices involving the processing of liquids at large pressure differences. High-speed photography gives a first insight into the bubble dynamics during the collapse [4],[5] and shows two fundamental phenomena during the non-spherical cavitation bubble collapse process: first the development of high-speed jets and second the release of shock-waves upon final bubble collapse. Both, the impact of shock waves and of high-speed jets on a surface can lead to material erosion. A more detailed experimental investigation including a precise determination of peak pressures at the wall and its association with the initial bubble configuration and evolution is beyond current experimental capabilities.

Cavitation is a major source of erosion for instance of ship propellers, pumps and water turbines. In such systems low pressure regions (pockets) exist where the water pressure suddenly becomes very low, almost a vacuum. These growing pockets, i.e. the “cavitation bubbles” propagate to high pressure regions, where they collapse immediately.

The deformation and breakup of a liquid drop in ambient liquid phase is a common phenomenon in a variety of scientific and engineering applications, such as oil pipeline transport, metallurgy, and certain types of emulsions when they encounter severe disturbances. Therefore it is of great importance to understand the behaviors and mechanisms of the drop under such circumstances.

For starting and running an engine successfully, the fuel must be mixed with highspeed air rapidly and fully in the combustor. The atomization quality of fuel impacts on the fuel mixture. Therefore study on fuel injection, atomization, diffusion and mixing quality is important, and themeasurement of fuel size and spatial distribution is the key to the engine research.

The use of sprays in liquid and gas flows covers a wide range of applications [1]. The efficiency of such multiphase systems to mitigate the devastating effects of blast waves which are issued from explosion is well known [2]. Indeed during shock loading, the water mist is subjected to aerodynamic forces from the carrier phase, which leads to a secondary atomization [3] of its individual liquid components. This induces a drastic increase of interfacial surface as well as interphase exchanges. Two-phase shock wave is known to show a long-time self-similarity behavior. The aim of this work is to investigate the transient pressure build-up due to secondary atomization and to confront the results from shock tube experiments with numerical modeling.

The study of particle cloud dispersion by a shock wave is important to many applications, including multiphase explosives that comprise a condensed explosive surrounded with packed micrometric reactive metal particles. Detonation of the explosive generates a shock wave which accelerates and compacts the particles. The particles are further accelerated from the reflected rarefaction as the shock wave arrives at the free-surface of the particles, leading to their rapid dispersal into the air. Generally speaking, the shock wave initially propagates through a granular particle bed and much later in time the particle cloud is dispersed and the flow becomes a dilute gas-solid flow. Between these two regimes the flow is characterized by a dense gas-solid flow [1]. Reactive multiphase flow models have been used to simulate the acceleration and dispersion of the particle cloud [2]. The fidelity of the simulations is severely limited by the physical drag models that need to take into account the interactions in the dense gas-solid flow [1]. In order to develop an accurate particle drag model that describes particle acceleration in the dense flow regime, representative experimental data need to be generated.While dispersal experiments have been carried out using a spherical multiphase explosive charge [3], controlled shock wave experiments are necessary to gain understanding of the fundamental physics of the dense supersonic gas-solid flow interactions and quantitative data concerning the particle cloud dispersion. Shock tube experiments have been carried out but typically the particle suspension method influences the shock flow [2] and the particle size is not typical of multiphase explosives. This paper reports on shock tube experiments looking at the acceleration and dispersion of 100 micron-sized aluminum oxide particles. The shock wave propagating into the packed particles provides a low-pressure analogy to multiphase explosive dispersal.

This work has been carried out to improve the methodology and understanding of mitigating against the effects of large explosive charges. This research specifically focuses on the physical processes involved in the expansion of granular materials. This is an area which has been investigated empirically in the past [1], but still lacks a thorough understanding. Hydrocode calculations are unable to predict the inhomogeneous expansion seen in experiments [2] and thus little confidence can be held in their ability to accurately predict the mitigating mechanisms.

Understanding the particle-particle and shock-particle interactions that occur in dense gas-solid flows is limited by a lack of knowledge of the underlying phenomena. Gas-solid flows are characterized by the particle volume fraction.

ϕ

p

of the flow [1]. For particle volume fractions less than about 0.1%, flow is considered dilute and the effects of particle collisions are negligible [2]. For packed particles, where the .p is greater than about 50%, the flow regime is said to be granular. The dilute and granular regimes have been well studied, but conversely, a substantial knowledge gap exists for dense gas-solid flows, which have intermediate particle volume fractions of about 0.1 to 50%. This regime exists at microsecond time scales during blast-induced dispersal of material when the shocked particles are closely spaced.

Gas-liquid interfaces, when subject to accelerations and/or velocity gradients, are unstable to infinitesimal perturbations—a classical subject that is well understood, mainly via linear theory, but individually for each class of such flows (Rayleigh- Taylor, Kelvin-Helmholtz, and Richtmeyer-Meshkov). Approximate, analytical, weakly non-linear and even non-linear methods exist for some cases, but again only for the rather idealized problems that involve accelerations normal to the interface

or

velocity gradients in flows parallel to the interface. While these inform qualitatively about systems found in practice, absent are understanding and capability to treat superposition of mechanisms in arbitrary flows, as for example those arising in the presence of accelerating, oblique or curved interfaces.More severely, absent are such methods that can accommodate compressible and shock-wave-bearing flows. This is the subject addressed by the numerical work summarized in this paper—the supporting experiments were carried out in a large-scale shock tube, they include Newtonian as well as viscoelastic liquids, and the quantification includes the resulting particle size distributions [1]. The canonical problem is aerobreakup of liquid drops [1], and applications of significant current interest include de-icing of airplane wings, internal-combustion, rocket, and pulse-detonation engines, and dissemination of liquid agents in the atmosphere.

Flow separation in convergent divergent (CD) nozzles occurs when the dynamic pressure of the fluid in the boundary layer is not high enough to overcome the rising pressure ratio along the diverging part of the nozzle. This phenomenon hasmany engineering application aspects such as jet engine performance, noise from jet engines, limitations on the construction of rocket nozzles or thrust vectoring. Therefore it is of interest to understand the physical aspects of the phenomenon.

Inside a nozzle with small opening angle the re-compression from supersonic to subsonic flow conditions is conducted by a shock train when the back pressure and the relative thickness of the boundary layer compared to the nozzle height are sufficiently high. A sketch of a typical shock train is depicted in Fig. 1a. Due to the boundary layer effect a series of shocks alternating with expansion zones is formed and the re-compression spans over a certain distance. The shock region is followed by a turbulent mixing region, where the subsonic turbulence layers originating at the foot of the first shock grow together. The turbulent fluctuations homogenise the flow and are accompanied by a rise in static pressure (see Fig. 1d). Caused by shear layers and the associated mixing process, the overall pressure rise in a shock train is lower than for a single normal shock with the same pre-shock conditions. A detailed description of the shock train phenomenon and a comprehensive review of its timeaveraged behaviour can be found in [1]. The normal Reynolds stress throughout the shock train on the centre line was measured by [2]. For the cases of a lambda-foot shock train and a x-type shock train a strong intensification of the turbulent fluctuations was observed in the mixing region. The magnitude of the maximum normal Reynolds stress on the centre line strongly depends on the pre-shockMach number. In the present investigation movable hot-wire probes are used to measure the distribution of the turbulent fluctuations over the complete nozzle height at different positions inside the shock train and the following mixing region. This allows to capture the growth of the turbulence layer. Figure 1b shows a Schlieren image of the flow with the hot-wire probe positioned on the centre line at the beginning of the turbulent mixing layer. Schlieren images with a shorter illumination time show the turbulent structures emerging from the shock/boundary layer interaction (Fig. 1c).

The recompression of supersonic gas flow is a very common flow phenomenon in modern aerodynamics and occurs in a great number of applications for instance supersonic ramjet or scramjet inlets, internal diffusers and supersonic ejectors. Under certain conditions even one or more shocks can appear downstream of the first shock. This series of shocks is a so called shock train. In contrast to other shock systems the supersonic flow is decelerated at first through a shock system and followed by a mixing region as shown in Fig. 1. For the whole interaction region Crocco et al. [1] have coined the term pseudo-shock. The structure and length of the shock train depends very much on the so called confinement level which is the ratio of the boundary layer thickness

δ

to the half height of the nozzle h. This was investigated very thoroughly by Carroll [2] and Om et al. [3]. In case of a shock train, due to the occurrence of successive shocks the pressure recovery along the shock train extends and greatly deviates form the pressure gradient that would occur at a single normal shock, [4]. However, a short recompression region or at best a single normal shock leads to very high pressure and temperature gradients, which provides interesting opportunities for various gas dynamic applications. In order to reduce the shock train length and thereby increasing the pressure gradient the shock system must be exposed to a higher back pressure. However, under normal conditions this would lead to a relocation of the shock train farther upstream, because the boundary layer can only withstand a certain back pressure level. The ability of the boundary layer to overcome a strong back pressure is limited, because the flow velocity hence impulse drops to zero near the wall. In order to remove the parts of the boundary layer with small impulse a normal suction slot 1.5 mm wide is placed upstream of the shock train, see Fig. 1.

During ascent in low altitudes, space launchers experience high nozzle exit pressure leading to flow separation in the divergent. Two flow separation regimes have been identified: free shock separation (FFS) and restricted shock separation (RSS) regimes [1]. The transition between these two regimes with potential asymmetry of the flow could lead to high side loads which are of great practical importance. Meanwhile, modern rocket nozzles are optimized for maximum thrust during the entire ascent trajectory. In compliance with the launcher overall dimensions, the divergent part of these nozzles is truncated and the nozzle area-ratio are readjusted to satisfy the exit Mach number. Thus, the resulting nozzle contour (TOC contour) is characterized by a high angle of divergence at the throat compared to an ideal nozzle contour. As a consequence, an internal shock emanating from the nozzle throat region can interact with the shock waves system produced by flow separation at the end of the nozzle. Recently for the first time, the influence of the internal shock wave on the transition between free separation and restricted separation regimes has been questioned [2]. An experimental work has been carried out in the framework of a PhD thesis at ONERA, Meudon, France. This work has had both a theoretical part by analyzing shock waves interferences inside the divergent and an extensive experimental program during which measurements of both stationary and non stationary pressures were obtained.

Operation conditions of rocket engines at sea-level and at high altitudes are different. The considerable difference in value of pressure of ambient air leads to different regimes of outflow from the nozzle - from the overexpanded regime at ground level to the underexpanded at high altitude. Different regimes of outflow from the nozzle cause considerable changes in flow pattern behind the nozzle outlet cross-section. One of the problems is the presence of shock structure in outgoing jet that causes flow separation from the wall of nozzle. It is known that separation point has no stable position on the nozzle surface, but fluctuates during the time. Boundary layer separation leads to the development of an asymmetric developed unsteady flow separation, and hence to unstable in time side-loads on the nozzle. Although modern rocket nozzles have a high longitudinal strength, they are not designed for side loads of high magnitude. The problem of high side loads was encountered in engine testing Vulcan rocket Ariane-5 [1, 2, 3]. Studies have shown that at the expiration of overexpanded regime two patterns of the jet of a nozzle implement, so-called free (Free Shock Separation, FSS) and restricted (Restricted Shock Separation, RSS) separation, with the side loads appearing in both cases [4].

Flow configurations with jet/supersonic flow interactions are met in different applications. Among them are supersonic aircraft cooling systems, thrust vector control devices, combustion chambers of hypersonic vehicles (scramjets), etc. From physical standpoint, the resulting flow structure represents complex phenomenon with separated, recompression and jet induced bow shocks, contact discontinuities and recirculating zones included. The jet itself possesses complex underexpanded structure, comprising barrel shocks and Mach disks. Thus, simulation of these types of flows is challenging task for numerical approaches, because accurate prediction of every flow detail has to be provided.

The acceleration of a fluid in a Laval nozzle to supersonic speed leads to a significant decrease of the static temperature in the flow, and deceleration via a normal shock results in a sudden reheating. Both effects are used within the joint project PAK 75 (Deutsche Forschungsgemeinschaft DFG) for the homogeneous ignition of a precursor and, hence, for the production of gas phase synthesized nanoparticles with narrow size distribution. However, the shock boundary layer interaction at the desired shock position leads to the formation of a so-called pseudo-shock system. Thereby, the heating rate across the shock system is reduced and the homogeneity of the particle growth is negatively affected by the inhomogeneous downstream conditions. Such pseudo-shock systems have been investigated by many research groups and a comprehensive overview is given by [1].

The flowfield resulting from the transverse gas injection into the supersonic cross-flow is the problematic of many aerospace applications ranging from the scram-jet fuel injection to the reaction jets and fluidic thrust vectoring(FTV). The prominent case of FTV by the use of the secondary injection represents promisingly attractive and effective way of control for small aerospace vehicles. Main advantages of the FTV are light-weightiness, simplicity and potential efficiency of such system comparing to the conventional mechanical TVC and mechanical deflectors. [5] Elimination of heavy and robust actuators and their replacement with only the fast-opening valves leads to the very significant reduction in mass. Fast dynamic response (~500Hz) to the conventional (~30Hz) [8] , very small losses in specific impulse and thus thrust are promising efficiency benefits. The CNES ”Perseus” project which this study is part of and design concept have aim of incorporating fluidic TVC on the future ”micro” launcher. Some of the results from the ongoing investigation are presented in this article.

A fluid with a certain speed escaping through a hole into an open space forms a jet; and its relevant phenomena are observed in a variety of situations including our daily lives [1]. Scientists and engineers have studied the jet flow for decades since it is related to crucial problems in aerospace engineering [2, 3], natural sciences, and various industries.

Transverse jet injection into supersonic/hypersonic cross flow has been encountered in many engineering applications ranging from scramjet combustors and solid rocket motor or liquid engine thrust vector control systems to high speed flying vehicle reaction control jets. These applications all involve complex three dimensional flow patterns comprising separated regions, shock waves, shear layers and wakes in common. Owing to numerous applications and these complicated flow features, transverse injections over different geometries and various forebodies have been received significant amount of interest. Earlier studies were focused on wind tunnel experiments and the utilisation of conventional measurement techniques such as Schlieren/Shadowgraph photography, wall pressure and concentration measurements to better understand the jet interaction and penetration phenomena. These studies aimed to assess the effect of injection pressure ratio, location of injection and state of incoming boundary layer and type of injectant gas on jets in supersonic/hypersonic cross flow. Recent studies of missile/forebody applications involving reaction control jets by several researchers have investigated jet interaction phenomenon on various axisymmetric body configurations at supersonic/ hypersonic speeds [1, 2, 3, 4]. Their aim was to investigate control effectiveness of transverse/lateral jets on different missile body configurations.

In the case of breakdown waves in a long discharge tube, near the electrode where the potential gradient in the gas is greatest, small quantity of gas is ionized. Analysis of the spectrum of radiation emitted from electric breakdown of a gas reveals no Doppler shift, indicating that the ions have negligible motion. The large difference in mobilities of positive ions and electrons causes establishment of a space charge and consequently a space charge field. The electric field accelerates the free electrons until they aquire enough of energy for collisional ionization of the gas. Since the ionized gas is a conductor and it can not hold internal electric filed, the electric field which has its maximum value at the interface between the ionized gas and the neutral gas has to reduce to a negligible value at the trailing edge of the wave.

The purpose of the present work is to provide insight into the numerical computation of shock- and detonation-waves, with a particular focus on high levels of grid refinement. This addresses the possibility of achieving a grid-independent solution and resolving the zone resulting from the ignition delay in the case of a detonation wave. Both viscous and inviscid flow solutions for shock and detonation waves are compared, and planar flow computations are used to assist with the setup of a twodimensional flow through a duct, blocked by 1/9-th of the inflow area by a cube placed on the centre-line of the duct.

A detached shock wave is formed in front of a blunt object moving with supersonic velocities or an object placed in a supersonic flow. The distance fromthe shock front to the surface of the moving body is called the shock stand-off distance,

δ

, and is measured along the propagation axis when the object is a solid sphere.

One undesirable problem in open cavity flows is the existence of strong and discrete cavity tones, especially with supersonic incoming flows.[1-4] A sound pressure level of almost 160 dB is observed for a supersonic cavity flow at Mach 2.0.[5] The strong cavity tones possibly result in structural vibrations and fatigue, adverse effects on store separation, and undesirable noise. The mechanism driving the cavity tones need to be clarified. The cavity tones are driven by self-sustained oscillations between the shear-layer instability and acoustic disturbances, which is named a feedback-loop mechanism.[6] Despite the fact that the feedback-loop mechanism itself has been well established and accepted, the dependence of cavity noise on variations of cavity configurations and flow conditions are not well-understood.[7]

Unlike standard large eddy simulation (LES) (for a review of LES for incompressible and compressible turbulence refer e.g. to [18, 7]), implicit LES (ILES) does not require an explicitly computed sub-grid scale (SGS) closure, but rather employs an inherent, usually nonlinear, regularization mechanism due to the nonlinear truncation error of the convective-flux discretization scheme as implicit SGS model. As finite-volume discretizations imply a top-hat filtered solution, regularized finitevolume reconstruction schemes were among the first ILES approaches, such as the flux-corrected transport (FCT) method [4], the piecewise parabolic method (PPM) [5]. Although ILES is attractive due to its relative simplicity, numerical robustness and easy implementation, it often exhibits inferior performance to explicit LES [8] if the discretization scheme is not constructed properly. Some schemes, such as PPM, FCT, MUSCL [16] and WENO [3] methods, work reasonably well for ILES by being able to recover a Kolmogorov-range for high-Reynolds-number turbulence up to

k

max

/2, where

k

max

is the Nyquist wavenumber of the underlying grid [9, 10, 21]. These promising results have led to further efforts on the physically-consistent design of discretization schemes for ILES. Physical consistency implies the correct and resolution-independent reproduction of the subgrid-scale (SGS) energy transfer mechanism of isotropic turbulence. Based on this notion the adaptive local deconvolution method (ALDM) has been developed [1, 11]. Approaches for decreasing excessive model dissipation for the solenoidal velocity field include the low-Mach number switch of [22], and the dilatation switch and shock sensor of [15].

Hypersonic flow computations still suffer from anomalous solutions such as a “carbuncle phenomenon” [1-3]. We still lack an accepted explanation for those anomalies, and we feel there is no single cause, nor is there any single cure. In the present study, we take the viewpoint that the shock anomalies are partly caused by the lack of mathematical expression for internal shock structure by the governing equations, and that they can be examined by numerical experiments. Quirk [1] introduced a benchmark test for numerical schemes on their responses to the captured (fast) moving shock. In this test, the shock took all the possible locations within a cell but instantly passed through them. Roberts [4] chose a more slowly moving shock which took 50 time steps to travel a single cell, and discussed a post-shock numerical noise propagation. Kitamura et al. [3] dealt with a stationary shock located within a cell with an initial shock position parameter

ε

=0.0, 0.1, ..., 0.9, i.e., the shock was placed at one of 10 possible locations in a cell. Their study discovered that any flux functions including FVS by Van Leer [5] are prone to carbuncles, though some of those methods had been believed to be carbuncle-free accoring to Quirk [1]. In other words, Quirk’s test for a moving shock was not enough to examine robustness of a numerical flux for a stationary shock. Moreover, in spite of those findings, the origin of the carbuncle remained a mystery. The present study will pursue its clue by bridging the gap between the works explained above, i.e., by varying the shock propagation speed from 0 of Kitamura et al. to 6 in Quirk’s choice and beyond. Following our earlier work, different flux functions by Roe [6], Van Leer [5], Liou (AUSM

 + 

 − up [7]), and Shima and Kitamura (SLAU [8]) will be used since they have different degrees of robustness against the shock.

The rearward facing step with a sharp corner is a classical configuration for studying separated flows. This configuration is also of practical relevance in a hypersonic vehicle. However, a truly sharp corner is basically a mathematical simplicity [9]. The influence of finite radius at the corner is therefore of significance and is studied here numerically under high enthalpy conditions using state-of-the-art computational fluid dynamics (CFD). The flow conditions are comparable to reentry velocity of 6.7 km/s which corresponds to a total specific enthalpy of

h

o

 ≈ 26 MJ/kg with a unit Reynolds number 1.82×10

6

per meter and a Mach number

M

 ∞ 

 ≈ 7.6 with air as the test gas. The geometry consists of an upstream flat-plate of length (

L

) of 48.4 mm with sharp leading edge. This is followed by a step of height (

h

) 2 mm and then a downstream flat plate of length (

D

) of 109.4 mm. Three different radii,

r

/

h

 = 1/8,

r

/

h

 = 1/3 and

r

/

h

 = 1/1 are considered here for the corner radius. The two-dimensional flow-field of interest is modelled using a time-dependent Multi-Block Compressible Navier-Stokes (MB-CNS) solver [6]. Perfect gas calculations were made with air to behave as a single species and real gas calculations were made assuming air to be a mixture of thermally perfect gas with 5 neutral species and adopting Gupta’s kinetic scheme for chemical reactions [3]. A multi-block structured grid with 52,000 cells is used. This was arrived at after performing a grid independence study over a sharp corner and modifying the grid topology to suit the curvature for the rounded configurations. Details regarding grid convergence are given in Deepak [1].

Explicit numerical schemes are widely used to simulate essentially unsteady flows with shock waves (e.g., see [1, 2] and numerous references there) because the use of large time steps with implicit schemes is often not possible and necessary due to time accuracy requirements. However, for some flows the time step of explicit time marching becomes severely limited by particular conditions within a relatively small flow area, as compared to the rest of the computational domain where the stability condition admits much higher time steps. The situation can be termed as “temporally-stiff”.Out of many examples, we mention the simulations of blast wave propagation when a high pressure/ temperature balloon is used as a blast wave source. When the blast wave propagates away from its origin, a high-temperature (and hence, high speed of sound) spot remains at the explosion center, considerably reducing the allowable time step for the whole simulation. The same effect can be caused by high flow velocities existing just downstream of a sharp corner when a shock wave diffracts over it, or by small computational cells near some small-scale geometrical feature of the problem under study.

The reduction of overall drag for an aircraft operating at its cruise condition is a prevalent goal for many involved in the aerodynamic design of transonic aircraft. Bringing both financial and environmental benefits, savings on the order of single drag counts are considered to be significant. Many concepts have been trailed in the pursuit of increased lift to drag ratios at transonic cruise. One of them, in terms of novel concepts addressing this issue, is the preservation of natural laminar flow over an increased proportion of the wing area. The laminarisation of a wing reduces overall drag via both a reduction in skin friction drag resulting from decreased mixing across the boundary layer and a reduction in pressure drag resulting from variations in effective aerofoil shape due to reductions in the boundary layer’s displacement thickness. Another concept which reduces the wave drag via the reduction of the strength of the near normal shock-waves on the wing upper surface is the application of 2D shock control bumps to the laminar or turbulent wings by Ashill

et al.

[1, 2]. Since Qin

et al.

[3] proposed 3D bumps for shock control, detailed experimental and numerical studies, including bump shape optimisation have been conducted by Wong

et al.

[4], Ogawa

et al.

[5] and Qin

et al.

[6]. While the

λ

–shock structure is observed as the key feature for some bump geometries tested [4, 5], e.g. ramp bumps, the optimised 2D and 3D bumps show a “knee-shaped” shock structure for smoothly (continuity of the first derivative) contoured bump designs [6].

Sectors of the aerospace and energy industries are amongst those interested in the efficient computational prediction of supersonic flow for both internal and external flow applications; e.g. the internal flow through engines and intake ducts, flow through nozzles, jet thrust vectoring. Shockwaves pose a numerical challenge due to the associated steep gradients in the flow field. Further physical modelling challenges arise from the interactions of these shocks with turbulent boundary layers and separated flow regions; so called Shock Boundary Layer Interactions (SWBLI). The high Reynolds numbers of many such applications mean that industry generally employs Reynolds Averaged Navier-Stokes (RANS) based approaches for turbulence.

Numerical modeling of flows with strong shock waves requires the use of fine meshes for an appropriate resolution. However, using high number of uniform grid cells might be time-consuming and require a large amount of computer resources. One of the way to weaken resources demands is using an adaptive mesh refinement (AMR) technique which tracks gradient of parameters in the cells [1] . A mesh is refined in the area of high gradients of solution, and a coarser grid is used in the case of low gradients.

Scramjet-powered access to space is expected to entail flight between Mach 5 and 15, along a dynamic pressure ascent trajectory of up to 2,000 psf (96 kPa) [1]. Scramjet engines typically need to be tested at sub-scale, even in the largest expansion tube facilities, and in these cases pressure-length (

p

-

L

) scaling is applied to maintain similarity for many flight parameters, including Reynolds number and binary reaction rates [2].

The development of high–speed flight and space access vehicles requires the solution of many technical challenges associated with the comparatively small net thrust at supersonic or hypersonic flight speeds. One of the more essential issues is the design of an air–breathing propulsion system capable of operating over the wide range of Mach (Ma) numbers, desired to facilitate the advancement of high–speed flight and space access vehicles. At flight speeds above Ma≈3 turbofan engines fall short since the compressed air through the engine reaches such temperatures that the compressor stage fan blades begin to fail. Instead ramjet engines, in which the profile of the air intake guarantees that the supersonic approach flow is decelerated to a subsonic flow through the combustor, where fuel is injected prior to mixing, self-ignition and combustion, may be used. However, beyond Ma≈5 extreme temperatures and pressure losses occur when decelerating the supersonic airflow to subsonic conditions, making the ramjet unpractical at higher flight speeds. At flight speeds beyond Ma≈5, supersonic combustion ramjets, or scramjets, in which the flow trough the inlet and combustor remain supersonic may be used. Achieving high combustion efficiency under such conditions, with residence time on the order of 1 ms, places extreme demands on the inlet, combustor, fuel–injector as well as on the nozzle design, [1]. The mixing of fuel and air, the self–ignition and the flame stabilization are thus critical processes.

Detonation engines are considered to potentially yield better performance than existing turbo-engines in terms of improved thermodynamic efficiency, simplicity of manufacture and operation, and high thrust-to-weight or thrust-to-volume ratio, amongst other advantages. Much effort has been put into the development of pulsed detonation engines (PDEs), including thermodynamic cycle analysis. Thermodynamic analysis of PDEs usually makes use of one-dimensional models, based on the Chapman–Jouguet (CJ) and the Zeldovich–von Neumann–Döring (ZND) theories, although increasingly sophisticated techniques, some involving numerical modeling, have also been developed lately. It is now understood that the Humphrey cycle used to model an isochoric cycle underpredicts the performance of a PDE [1]–[4]. The so-called Fickett–Jacobs (FJ) cycle is based on the CJ model. While an improvement over the Humphrey cycle, its reliance on the CJ model means that it fails to account for the physics espoused by the ZND model [1, 2]. In this paper, a discussion of the Humphrey and FJ cycles is given and the proper ZND cycle is suggested. These cycles are illustrated with a hydrogen/air mixture initially at STP. The use of a generic heat release parameter to construct the ZND cycle is provided.

Today the access to space is done, only, by multi-stage rocket-powered vehicles, which have flown hypersonically, carrying their own propellant (solid and/or liquid, oxidizer along with fuel) to propel payloads and astronauts to Earth’s orbit.

The modern aerospace vehicles utilize multi-stage propulsion system on board, in general not reusable, of combustion chemistry (solid propulsion and / or liquid propulsion), extracting and converting chemical energy into kinetic energy with 97-98% efficiency. Approximately 89% of the weight of the spacecraft at time of launch, is due to the propulsion system be part of the vehicle , with only 1 to 2.5% due to the payload , usually satellites.

A new generation of scientific aerospace vehicles, using advanced hypersonic airbreathing propulsion based on supersonic combustion technology, is in development at several research centers [1]. This new propulsion system (scramjets) is economically and ecologically more attractive than the conventional rocket propulsion.

Metal particle contamination is a concern for liquid rocket engines that use enriched O

2

at high pressure. It is believed that under some engine conditions contaminant particle impact could release sufficient kinetic energy to initiate combustion, providing an ignition source for engine components (e.g., turbine blades) impacted by the particles, and subsequently a combustion event that eventually consumes structural materials of the engine. It is important that the combustion properties of these candidate metal particles be studied for their propensity to cause ignition under rocket-like conditions, to reduce the risk of engine failure. Laboratory study of such a mechanism under realistic engine conditions is difficult, and data are lacking. Data that reveal the influence of particle mass, kinetic energy, impacted-surface composition, and environmental conditions on ignition propensity are valuable for launch programs involving oxidizer-rich, staged combustion engines.

The air inlet is a crucial component of hypersonic airbreathing engines, which should decelerate and compress airflow with minimum losses. For efficient engine operation the inlet must be started, i.e., all incoming supersonic flow must be captured and the flow inside the inlet must be predominantly supersonic. Kantrowitz and Donaldson [1, 2] established the classical theory of flow starting in converging ducts. According to the theory, the limiting duct area ratio for spontaneous starting (or self-starting) is based on the flow condition at which a normal shock is positioned exactly at the duct entry and the post-shock subsonic flow isentropically accelerates along the duct to become sonic at the duct exit (i.e., the choked throat is considered). It is assumed that the flow is quasi-one-dimensional and quasi-steady. For exit-toentry area ratios exceeding the limiting values, which depend on freestream Mach number, the duct (inlet) flow would start on its own, upon the increase of freestream velocity fromzero to the required value. As follows fromthe Kantrowitz theory, limiting contractions for starting lead to low contraction inlets, which do not provide sufficient compression for scramjet operation. Practical, high-contraction inlets do not satisfy the Kantrowitz self-starting condition and would not start spontaneously. This constitutes a well-known inlet starting problem.

Interaction between shock wave and combustion is very important for supersonic combustion. For scramjet, isolator is a key element to withstand the high pressure due to combustion and to avoid the unstart of the inlet. Therefore, the flow is very complex in isolator and combustor because of the interaction between combustion and shock wave. Usually, there are two modes of combustion in scramjet: supersonic mode and subsonic mode. Many researches have already shown how to achieve dual-mode scramjet to obtain better engine performance [1] [2] [3].

However, the mechanism of dual-mode combustion is still unclear. In this paper, experimental and numerical investigations were attempted for better understanding of the dual-mode combustion for scramjet applications.

Often, and especially in canonical turbulence research, the belief is that initial conditions wash-out and the turbulence develops to a universal self-similar state [1, 2]. However, recent numerical work [3, 4] has shown that this hypothesis hold true only for some flows, and that the buoyancy driven (Rayleigh-Taylor) turbulence is dependent upon initial conditions, and a self-similar state has not been measured in experiments. Similarly, R-M flows, driven by a shock wave, have a time-dependent mixing evolution that is also dependent upon initial conditions [5, 6, 7, 8]. In this present study we focus on improving our understanding of the nature of initial conditions on R-M mixing.

When a shock wave impacts an interface between two fluids, the misalignment of the pressure and density gradients results in deposition of baroclinic vorticity on the interface. This leads to the growth of initial perturbations on the interface causing the phenomenon of Richtmyer-Meshkov instability (RMI) [1]. The RMI produces turbulent mixing of the fluids which plays an important role in many physical and technological processes like inertial confinement fusion, supersonic combustion, and impact dynamics of liquids. The RMI is also believed to be the reason for the increased mixing observed in the optical output of supernova 1987A. Understanding this process requires robust numerical algorithms capable of simulating this highly non-linear flow and high quality repeatble experimental data to validate the numerical findings. Experimental limitations (difficulty generating a well characterized initial interface between two fluids, and diagonostic limitations) and shortcomings of numerical algorithms have constrained detailed explorations of the physics of this intability. Recent improvements in experimental methods (like membrane-less techniques to generate an interface) have provided reliable data for numerical code validation and simulation of this flow to characterize the turbulent mixing between the fluids. In this work one such configuration of a shock in air impacting a curtain of cylinders of dense gas is chosen for numerical simulation. The results are compared to available experimental data [2] and some turbulence statistics are reported.

When the material interface separating two different fluids is accelerated by shock wave, a hydrodynamic instability happens, which is well known as the Richtmyer-Meshkov instability (RMI) [1,2]. The physical mechanism for the occurrence of RMI is the deposition of baroclinic vorticity produced by the misalignment of the pressure gradient of the shock wave and the local density gradient at the interface (i.e.

$\nabla\rho\times\nabla\rho\neq$

0). Another type of instability is called the Kelvin-Helmholts instability (KHI) [3], which is because of the presence of tangential velocity jump at the interface. At late times of the RMI developing, because of the larger velocity difference at both sides of the spike and at the tip of the bubble, the KHI also starts to develop. The RMI is of importance in a wide range from man-made applications to natural phenomena such as inertial confinement fusion (ICF) and astrophysics. The KHI also has a prominent significance in plasma flow, radioactively driven molecular clouds [4], etc. So they have gained much attention for many years.

When an initially perturbed interface separating two fluids with different properties is impulsively accelerated by a shock wave, the flow field will exhibit complex fluid dynamic phenomena due to the misalignment of the density and pressure gradients. It is often referred to as the Richtmyer-Meshkov (RM) instability [1, 2] and has been investigated within the past several decades due to its extensive physical applications such as inertial confinement fusion (ICF) [3], turbulent mixing in scram jet [4] and collapse in supernova [5]. Specifically, in most applications the shock waves maintain two-dimensional (e.g. cylindrical) or three-dimensional (e.g. spherical) converging shapes. Taking the ICF experiments, the thermonuclear fuel is contained in a small spherical solid capsule in advance and illuminated evenly by a number of well-designed laser beams. Simultaneously a spherical converging shock wave is generated and traverses the capsule. In the process even a tiny imperfection of the capsule surface can induce the RM instability and may destroy the fusion reaction [3]. It is therefore of fundamental interest and practical significance to explore the fluid dynamics of the flows in the interaction of interfaces with a converging shock wave.

Richtmyer-Meshkov (RM) instability occurs when a shock wave passes an interface that separates two media with different densities [1, 2]. Its research is of importance in inertial controlled fusion (ICF), shock-flame interaction in Scramjet engines, detonation wave generation and propagation in pulsed detonation engines, volcanic eruption, vapor explosion of nuclear fuel in nuclear power plants, etc. RM instability also appears in supernova explosion in astronomy and it has often been used in modeling the formation of fixed stars. On the other hand, since turbulent mixing becomes dominant in the later stage of RM instability, its study is theoretically meaningful in understanding turbulence problems [3]-[6]. The first theoretical model of RM instability was given by Richtmyer in 1960 [1]. He proposed an impulse model considering fluid compressibility. The first experiment of RM instability was done by Meshkov in 1969 [2]. Later, Benjamin and Fritz [7] conducted experiments of RM instability on a shocked interface between liquids having a density ratio of 10. In 1972, Myer and Blewett [8] simulated RM instability using Lagrange method and their results are qualitatively in agreement with that of Meshkov’s experiment. Recent investigations have shown that the early stage of the instability is compressible and nearly linear and its later stage is nearly incompressible and nonlinear. In fluid mechanics, RM instability is a typical and difficult problem whereas innovation in experimental techniques is necessary in research deeply into RM instability phenomenon. Due to the great density difference between a gas and a liquid, it is convenient to visualize a gas-liquid interface in RM instability [9]-[12]. Based on these work, we put a layer of silicon oil on the water column. Thus, the oil layer is bounded by a gas-oil interface and an oil-water interface. Therefore, when a shock wave passes through the layer, RM instabilities with the Atwood number

A

t

 = 1 and

A

t

 = 0 all occur simultaneously. This means that an interface with a wide range of Atwood number from 1 to 0 has been constructed and the experimental capability of the facility has been extended. The definition of the Atwood number is

$$ A=\frac{\rho_2-\rho_1}{\rho_2+\rho_1} ~~(1) $$

where

ρ

1

and

ρ

2

are the fluids densities at two sides of the interface respectively. This paper’s results have not been seen in open literatures.

A planar shock wave that impulsively accelerates a spherical density inhomogeneity baroclinically deposits vorticity and enhances the mixing between the two fluids resulting in a complex, turbulent flow field. This is known as the classical shockbubble interaction (SBI) and has been a topic of study for several decades [1,2,3,4, 5,6,7,8,9,10,11,12], and closely related the Richtmyer-Meshkov instability (RMI) [13, 14]. While the classical SBI problem concerns a reactively neutral bubble, the present experimental study is the first of its kind in which a spherical bubble filled with a stoichiometric mixture of H

2

and O

2

diluted with Xe is accelerated by a planar shock wave (1.35 < M < 2.85) in ambient N

2

, and will be referred to as reactive shock-bubble interaction (RSBI).

The shock-accelerated inhomogeneous flows have been widely investigated for the fundamental interests and diverse applications in a broad range of spatial, temporal and energy scales, such as the supernova explosions, supersonic combustions and inertial confinement fusion (ICF) implosions. The particularly simple configuration, the shock-bubble interaction [1], has been considered as a basic configuration to study the flows. With respect to the development of the density-stratified interface impulsively accelerated by a shock wave, analogies may be drawn to the study of the Richtmyer-Meshkov (RM) instability [2, 3].

Richtmyer-Meshkov (RM) instability occurs when a interface separating two fluids of different density is impulsively accelerated in the direction of its normal. It is one of the most fundamental fluid instabilities and is of importance to the fields of astrophysics and inertial confinement fusion. RM instability experiments are normally carried out in shock tubes, where the generation of a sharp, well-controlled interface between gases is difficult, so there is a dispersion in terms of experimental results. The experiments presented here were conducted in a horizontal shock tube where the materialization of the initial interface was achieved by a thin nitrocellulosic membrane (0.5

μ

m thick) deposited on a stereolithographed grid support, computer-aided designed and constructed with chosen shape and dimensions. As diagnostic, we used laser sheet flow visualization to yield time-motion image sequences of the linear and the non-linear developments of the instability. In previous investigation [1], we have already shown that residual pieces from the membrane constituting the initial interface tend to delay the interpenetration in the light-to-heavy gas configuration and specifically during the linear stage of the interface evolution. In order to reduce these effects in the present experiments, we have increased the strength of the shock wave (Mach~1.5). We have also extended the test section from 0.46 m to 1.5 m which allows the instability to grow further and thus to observe the whole nonlinear phase until the transition to turbulence. The present paper summarizes the results obtained during this study undertaken for air/SF

6

and air/He gas combinations (positive and negative Atwood numbers, respectively) in 2D and 3D geometries.

The Richtmyer-Meshkov instability (RMI) develops when a shock wave traverses a density interface separating two gases. The miss-alignment of the pressure gradient across the shock and the local density gradient at the contact during shock passage leads to vorticity production at the interface. Subsequently the flow driven by the deposited vorticity leads to interfacial instability growth and eventually to turbulence mixing. RMI is important in many areas of physics, from geophysical and astrophysical problems to industrial applications. In particular, attention has recently focused on RMI and RM mixing in the converging geometry such as that occurs in an imploding inertial confinement fusion (ICF) capsule. When a stratified cylindrical shell with initial perturbations is driven by a convergent shock wave, the effect of convergence tends to enhance the perturbation growth compared with that in a planar geometry. The convergent incident shock wave reflects at the cylinder center and the succedent reflected shock waves move to and fro within the whole region. Besides, the Rayleigh-Taylor instability (RTI) also occurs whenever the light fluid accelerates the heavy one during the evolution. All these factors make the mixing procedure in a stratified cylindrical shell driven by shock wave much complex than that in the planar geometry. Although many models have been proposed to predict the instability growth in the linear, weakly nonlinear, and turbulent regimes, each of these models has limitations and a restricted domain of applicability. For this complex mixing process with strongly nonlinear transition stage, the direct numerical simulation with high resolution is the common way to study its evolution. In this paper a hybrid scheme combined with the finite-difference and the weighted essentially non-oscillatory (WENO) method, combined with high order strong-stability preserving Runge-Kutta scheme for the time integration, is used to simulate the mixing due to the interfacial instability in a stratified cylindrical shell driven by convergent shock wave. Growth and mixing properties of the turbulent mixing zone (TMZ) have been investigated using the simulation results. And the effect of initial perturbation on the mixing has been discussed.

Lombardini et al. [1] have recently carried out Richtmyer-Meshkov instability (RMI) simulations in cylindrical geometry. RMI for a spherical axisymmetric flow was investigated by Dutta et al. [2].We consider amore general initial interface perturbation, with a spherical egg-carton profile similar to the one used in planar RMI simulations [3, 4]. An interesting feature of this profile is that the perturbationwavelength is nearly constant over the spherical shell spanned by the material interface. The fluids considered in this study are air outside and SF

6

inside. The shock is launched from the air (lighter) side of the interface. As the flow evolves, a series of reflected and transmitted shocks are generated,which via baroclinic deposition of vorticity and its subsequent transport serve to mix the two fluids in a turbulent mixing zone.

The Richtmyer-Meshkov Instability (RMI) occurs in several physical and technological processes such as supernova explosion, supersonic combustion, detonics or inertial confinement fusion. This instability develops when interfacial perturbations, between two fluids of different densities, grow because of a shock wave induced impulsive acceleration. The basic mechanism for the initial growth of perturbations on the interface is the baroclinic generation of vorticity which results from the misalignment of the pressure and density gradientswhen the shock crosses the interface. Early time linear and following nonlinear growth of the RMI have been, and are still widely investigated, either theoretically, numerically and experimentally [1]. Nevertheless, experimental investigation of the Turbulent Mixing Zone (TMZ) induced by a rapidly growing RMI is still nowadays poorly documented, even if we can mention for instance the work of Leinov et al. [2] who characterized the growth of the MZ with time following the passage of the re-shock (with an emphasis on the influence of the initial amplitude of the MZ and the reshock strength), and the study of Poggi et al. [3] in which the production of turbulence by the TMZ has been investigated in a vertical shock tube using two-components Laser Doppler Velocimetry (LDV).

The main objective of the present work is to provide a detailed characterization of our shock tube in order to discriminate the turbulence level produced by the mixing of the two gases in the TMZ, through baroclinic effects, from the background turbulence level of the experimental set-up. We thus investigated several configurations without the mixing zone (pure air).

Membraneless Richtmyer-Meshkov instability experiments have previously been carried out in a vertical shock tube using a single-mode two- and three-dimensional initial perturbations [1], [2]. The present study utilizes the apparatus and experimental techniques of these previous investigations modified to allow the generation of a random three-dimensional initial perturbation.

A 5m long vertical shock tube with a 10.2cm diameter round driver, and a 8.9cm square test section is used for this study. The light gas (air) enters the tube at the top of the driven section immediately below the diaphragm, and the heavy gas (SF

6

) enters at the bottom of the test section. The gases exit the shock tube through a series of small holes in the test section walls, leaving behind a flat, diffuse interface. In the previous studies the initial perturbation was generated by periodically oscillating the square shock tube, laterally, to produce a nearly single-mode two-dimensional standing wave. More recently [3] we have found that we can produce similar single mode three-dimensional standing waves by oscillating the gas column within the shock tube vertically using the periodic motion of a piston mounted at the bottom of the test section. The work presented here is a continuation of that work in which the frequency of this motion is increased producing a more random, short wavelength pattern.

Micro vortex generators are a new kind of passive flow control instruments for shock-boundary layer interaction problems. In contrary to the conventional vortex generator, they have heights approximately 20-40% (more or less) of the boundary layer. Among them, Mircoramp vortex generators (MVG) are given special interest by engineers because of their structural robustness. The mechanism of the flow control was thought that a pair of streamwise vortex is generated by MVG and remains in the boundary layer for relatively long distance; the down-wash effect by the streamwise vortices will bring about momentum exchange, which makes the boundary layer less liable to separation. During such process, a specific phenomenon called as momentum deficit will happen [1], i.e., a cylindrical region consisted of low speed flows will be formed after the MVG. It was pointed out by Li and Liu [2] that the origin of deficit comes from the shedding of boundary layer over MVG.

Numerical simulations have been made on MVG for comparative study and further design purposes. Ghosh, Choi and Edwards [3] made detailed computations under the experimental conditions given by Babinsky by using RANS, hybrid RANS/LES and immersed boundary (IB) techniques. Lee et al [4, 5] also made computations on the micro VGs problems by using Monotone Integrated Large Eddy Simulations (MILES). Basic flow structures like momentum deficit and streamwise vortices were reproduced in the computation. Further studies were also conducted on the improvement of the control effect.

Shock-wave boundary-layer interactions (SWBLI) are prevalent in many supersonic applications, e.g., over deflected flaps, fore-body ramp corners, on leading edges where the bow shock from the vehicle nose interferes, along axial compression corners inside air-inlets, shock reflection and crossing-shock interactions in the inlets etc. The adverse pressure gradient across the interaction shock can cause separation of the incoming boundary-layer leading to increased aerodynamic drag, heat transfer and unsteady pressure loads. Much of the early work over forward-facing steps [1], un-swept compression ramp flows [2-4] and in interactions induced by blunt fins [5], circular cylinders and sharp fins at angle of attack [6] was focused on understanding the dynamic/unsteady behavior of these interactions. It has been observed that the flow in these interactions in unsteady if the pressure ratio across the oblique shock is such that the mass of the fluid reversed at the reattachment point does not balance the scavenged fluid from the separated region [7-8]. As a result, the separated region “breathes” and during one half of pulse, mass is injected into it while during the other half it is ejected out resulting into an unsteady mass exchange.

Shock wave boundary layer interaction (SWBLI) is a flow phenomenon that is critical for many high speed applications, such as supersonic inlets and propulsion-wing or -fuselage interactions. Much effort has been devoted to investigate the mechanism of SWBLI, its turbulent nature and the role of large-scale fluctuations[1]. Different types of flow control techniques have been proposed to alleviate the adverse effects introduced by SWBLI, such as flow separation and fluctuating pressure loads. Micro-ramp vortex generators are considered to be a preferred type of passive boundary layer control technique, due to a limited increase in drag compared to conventional larger vortex generators that emerge outside the boundary layer, while still being effective in reducing flow separation.

Three-dimensional crossing-shock-wave and turbulent boundary-layer interactions can generate intense wall heat flux rates, high pressure levels, and large-scale flow separations on high-speed vehicle surfaces. To reproduce such complex flow physics, simple configurations, such as single-sharp fin and double-sharp fin mounted on a flat plate, were adopted in previous investigations. Review papers by Knight

et

al

. [1] and Zheltovodov [2] provided summary of current state-of-the-art of research advancements in this field. Recently, Yao

et

al

. [3] carried out numerical simulation of symmetric double fin configurations of 7°×7°, 11°×11°, and 15°×15° wedge angles. Results of surface static pressure distributions were found in good agreement with wind tunnel experiments [4] and other numerical simulations [5] but heat flux coefficient distribution differed from experimental data at the 15°×15° case. In this work, an additional configuration of 19°×19° case is introduced to investigate flow topology due to increased shock-viscous interaction strength. The predicted flow field will be compared qualitatively with available experiments [6, 7] and other relevant numerical studies [7, 8].

In the literature, we can find lot of analytical or numerical studies about shock wave interaction. However, in the major part of this work, the two dimensional assumption is used [1]-[4]. Although we know that in real flight conditions the interaction is at least axisymmetrical or three dimensional, we have also chosen to deal with axisymmetrical interaction. Indeed, the comprehension of the axisymmetrical phenomena is needed before taken into account a more realistic three dimensional case. Thus, in the present paper, we have numerically studied the interaction between a shock generated by a conical ring and a shock generated by a sphere (respectively called the conical and spherical shock). A schematic description of the study case is given by the figure 1. The inlet conditions are:

M

=4.96,

T

=77

K

and P=1700

Pa

. The preliminary results on this topic have been presented in [5] and [6].

Interactions of shock-waves with boundary layers are a common feature in highspeed flight. Depending on the nature of the incoming boundary layer such interactions may lead to large unsteady thermal and pressure loads which may reduce the aerodynamic performance and the structural integrity of hypersonic vehicles. Despite numerous investigations our current knowledge of the fundamental physical mechanisms involved in unsteady shock-wave/boundary-layer interactions (SBLI) is far from complete and a number of aerospace applications would benefit from a deeper understanding of the subject. Most of the research efforts in this field have been directed to the analysis of shock-waves interacting with nominally twodimensional turbulent boundary layers [1]. Flows over high-speed vehicles and, in particular, inside the intakes of their air-breathing propulsion systems are very complex and include interactions of shock-waves with three-dimensional transitional boundary layers. The transition process is very sensitive to flow conditions and geometric parameters. Experiments have shown that small roughness elements, less than a millimetre in height, can lead to early breakdown to turbulence even in a quiet environment [2]. In high-speed flows, transitional boundary layers can also be affected by the interaction with shock-waves through mechanisms which are largely unknown. A detailed study of three-dimensional transitional SBLI will help understand how shock-waves affect the transition process at high-speeds. The limited number of studies available in the literature on transitional SBLI show that for strong interactions (in the convective instability regime) small-amplitude disturbances experience strong amplification across the separation bubble due to the instability of the separated shear layer [3]. In addition, transitional interactions induce higher levels of unsteadiness and stronger thermal loads than in the fully turbulent case [4, 5].

The study of supersonic turbulent flows in the channelswith sudden expansion (step/ cavities) is actual task since this configuration is used for ignition and flame stabilization. Supersonic combustion is studied for many years in order to support the future hypersonic flights. It well known that it is rather difficult to get the ignition and stable combustion at supersonic speeds [1]. The flow in the supersonic combustion chamber js charachterized by a short residence time which is only a few milliseconds of magnitude. A simple geometry to generate a flameholding region in supersonic flow is a backward facing step (BFS). Flows around BFS configuration were studied for decades and many papers were published regarding fundamental flow properties [2] as well as the scramjet combustion chamber utilization [3]. The effect of step configuration (“boattailing”) on the structure of compressible base flows was investigated in a numerous papers since Hama’s work [4]. Nevertheless, the question is little studied as far as supersonic chemically reacting flows in channels are concerned.

Variable camber concept, such as deflection of flaps or ailerons in cruise, could play a role in performance optimization for current- and future-generation aircrafts. Within the operational flight envelopes, it would change the flowfield and consequently modify aerodynamic characteristics[1]. A study by Szodruch and Hilibig[2] also indicated that variable camber can be employed to improve the transonic maneuvering characteristics of a fighter aircraft. Furthermore, Parndtl-Meyer expansion is well known in supersonic flows. However, the subsonic expansion or transonic expansion flows around a sharp convex corner are less studied. At lower Mach number, the flow is expanded and recompressed around a sharp corner. With increasing Mach number or convex-corner angle, the boundary layer is subject to a rapid acceleration and the flow switches to transonic expansion flow. Noted that Chung[3] proposed a similarity parameter to characterize the flowfield, in which the transition of subsonic and transonic expansion flows is observed at

M

2

η

=6.14. Shock-induced boundary layer separation is also another concern for application of variable camber concept.

The characteristic of aircraft drag is one of main qualifications to judge the performance of an aircraft. Aircraft drag reduction is a very actual and important problem. One percent drag decrease lets, approximately, to 10 percent increase aircraft payload or increase long-range.Many reports[1-7]were made on this subject presenting multifarious new technique for drag reduction. Hereinto, it is a novel approach for reducing aerodynamic drag by employing plasma.

Shock wave/boundary-layer interaction (SWBLI) is a common but important flow phenomenon, within various engineering design areas such as engine inlets, compressors and turbines. In past decades, researchers have made great efforts towards better understanding and modeling of SWBLI flows. The reviews by Knight and Degrez[1], Zheltovodov[2], Dolling[3] examine the capability of Reynoldsaveraged Navier-Stokes (RANS) turbulence models in the prediction of SWBLI. The common conclusion is that most RANS models based on the linear formulation of the Boussinesq assumption are difficult to accurately predict details of flow separation, i.e. the distributions of pressure loads, heat transfer and skin friction. In consequence, a lot of efforts are put in deriving non-linear RANS turbulence models, either in an explicit algebraic form or through transport equations for the Reynolds stress components. The ongoing research emphasis is to get a physically reliable understanding of SWBLI and to reach a point where a unique non-linear formulation could be used for the modeling of Reynolds stress in a large range of flow configurations.

In ideal shock tube experiments, flow properties behind the incident and reflected shock waves do not vary with distance or time, and can be calculated using the standard normal shock equations and the known incident shock speed. However, nonideal effects result in flow nonuniformity behind the incident shock wave, leading in turn to nonuniformity behind the reflected shock. It has been observed that behind the reflected shock, pressure typically increases gradually with time (positive dP5/dt) [1-3]. Such deviations of pressure (and concomitantly the temperature) from ideal values cause errors in chemical kinetic studies of rate coefficients and ignition delay times [2, 3]. Thus it is worthwhile to analyze nonideal shock tube effects and determine their impact on flow conditions, with a goal of improving experimental methods and gasdynamic models that will lead to more accurate experiments.

The main function of a scramjet inlet is to capture air flow from the incoming hypersonic stream, compress it through a series of shocks or compression waves, and provide uniform flow to the combustor. There should be maximum mass capture along with a minimum stagnation pressure loss in the inlet.

The shock waves in the inlet duct interact with the boundary layer on the walls and can result in flow separation due to strong adverse pressure gradient across the shock wave. The shock/boundary-layer interaction often results in a complex flow pattern, comprising of additional shocks, expansion waves, shear layer and separation bubble. The separation bubbles are highly viscous, and hence increase the stagnation pressure loss. Peak values of pressure, skin friction and heat transfer rates are found at reattachment point. Also, the separation bubble acts as a blockage to the flow inside the inlet duct and can result in inlet unstart. It is therefore important to predict the shock/boundary-layer interactions in a scramjet inlet, including the size of the recirculation region, accurately.

Oblique shock reflection phenomena have been long considered to preserve selfsimilarity. However, the authors’ investigation of von Neumann reflection [1] was a turning point that cast doubt on self-similarity. The von Neumann reflection is a new type of oblique reflection first referred to by Colella and Henderson [2]. This reflection is geometrically characteristic in that a Mach stem is tangentially connected with an incident shock at a triple point. Furthermore, its reflected wave is very weak and a slipstream is barely optically observable (Fig. 1 (a)), and stands in contrast to ordinary Mach reflection (Fig. 1 (b)). This reflection takes place when the incident shock Mach number

M

i

and/or the reflecting wedge angle

θ

w

are small. According to Colella and Henderson [2], von Neumann reflection exhibits the von Neumann paradox. For weak shock waves, Kobayashi et al. [1, 3] found that, as the incident shock proceeds, the wave angles vary along a trivial solution curve of von Neumann’s three-shock theory, while the triple point moves along a straight line passing through the wedge apex as if self-similarity holds. Further investigations [4] revealed experimentally that oblique shock reflection in a shock tube is generally non-self-similar, even though the triple-point trajectory is linear. This non-selfsimilarity is a new idea requring reconsideration of the von Neumann paradox.

The physical understanding and modeling of shock mitigation are important for the development of an effective barrier arrangement related to disaster management. While it is not currently feasible to simulate and analyze full configurations in detail, sufficient progress has been made to analyze the dynamics of simpler building block flows that provide useful insights into the underlying dynamics of these complex flows. Also, apart form the experimental study, numerical simulation has become quintessential tool for prediction of complex physics in solid/fluid interaction problems. Several authors dealt with experimental or numerical approaches in order to study the unsteady shock wave interaction with multiple obstacles, such as cylinders, spheres and triangular prisms [1, 2, 3, 4]. According to the recent findings of [5], the influence of different geometrical shapes on shock-wave attenuation is negligible for higher open passage. However, this finding requires a systematic study of the effects of different parameters for lower values of the open passage. In our previous works [6, 7], excellent agreement between experimental and numerical results is obtained for the case of shock-wave interaction with single cylinder and triangular prism. These validations prove the reliability of the computational techniques used for the present study. It is being observed that after the passage of the shock through the obstacle matrix, eddies of different length scales are generated, but the later stage of shock-vortex, shocklet-vortexlet interaction are different for inviscid and viscous computations [8].

To resolve the von Neumann paradox, Guderley (1947) proposed a four-wave structure, consisting of an expansion fan in addition to the three shock wave configuration [1]. Despite intensive experimental work at the time, no evidence of this expansion fan was observed, therefore dismissing Guderley’s proposal. Experimental work conducted by Skews and Ashworth [2] with the use of a large-scale shock tube showed experimentally for the first time that Guderley’s proposal was, in fact, correct.

Upstream moving pressure waves are observed in transonic flows over airfoils already for decades. They can be generated actively by flap or airfoil oscillations [6]. But, they are also naturally present in airfoil flows [1, 4]. Upstream moving pressure waves can lead to flow instabilities by forcing shock and stagnation point oscillations. Furthermore, it is expected that they affect the laminar-turbulent transition [5]. Hence, the phenomenon is of great engineering interest.

In spite of considerable research effort in past decades, the reflection of shock waves off convex cylindrical surfaces still poses a number of unanswered questions. For a given shock Mach number

M

S

, the reflection pattern changes from regular to irregular at a certain wall angle Θ

w

. If one determines this transition angle by visual inspection of the reflection pattern and defines it as the location of the first occurrence of a visible Mach stem, one typically arrives at wall angle values lower than the one found in the pseudo-steady case for a straight wedge at the same Mach number [1]. This would indicate that the regular reflection pattern is maintained longer on the cylindrical surface compared to the straight wedge case. Numerical simulations, on the other hand, suggest that the transition occurs at the same wall angle as for the straight wedge. If this were the case, the transition would be governed by the local wall angle and would not be influenced by the preceding history of the reflection. Furthermore, the delayed transition observed experimentally would then have to be explained as a consequence of insufficient spatial resolution of the optical records used to classify the type of reflection: If the Mach stem is too small to be optically resolved, the reflection pattern will remain to appear regular beyond the transition angle predicted by the theory for straight walls.

The aerodynamic characteristics of an object flying in close proximity to a solid surface are altered by the presence of this boundary. This phenomenon, known as aerodynamic ground effect, has been thoroughly investigated in the subsonic flow regime [1], but to the best of the authors’ knowledge, high-speed applications of this effect with objects flying at transonic speeds have only been briefly considered on a theoretical basis [2]. Furthermore, most of the ground effect literature concentrates on lift-generating objects such as wings and/or aircraft where the presence of the ground alters an existing but inherently asymmetric pressure distribution. A non-lifting body - a projectile - is arguably a simpler case as the ground effect would immediately become obvious through an asymmetry in the originally symmetric pressure distribution. This was clearly shown in an earlier study on projectiles at medium supersonic speeds (

M

= 2.4). In this case, the pressure distribution around the projectile remained unaffected by the presence of the ground until the reflected bow shock impinged on the projectile. From this point onward, the pressure distribution was drastically changed and led to a non-negligible lift force, an associated pitching moment and a change of the drag (the latter influence was primarily due to a change of base pressure). As soon as the reflected shock did no longer impinge on the projectile or the near wake, the pressure distribution was not affected and the aerodynamic ground effect became negligible and eventually non-existent. For the investigated projectile at

M

= 2.4 this occurred for clearances h/d > 1, where h is the clearance between the projectile and the ground and d is the projectile diameter. Experiments and numerical simulations established the trends for lift, drag and aerodynamic moment [3].

As it can be used to efficiently concentrate energy through high compressions [1], the collapse of imploding shock waves is of great interest. The stability and the modeling of these waves are still subject to studies. Furthermore, when a shock wave crosses an interface between two materials, this interface becomes unstable and the Richtmyer-Meshkov (RM) instability develops [2, 3]. In spherical geometry, experiments about this instability are scarce. As a consequence, there is a strong need of experiments about spherically imploding shock waves.

S. D’yakov was the first who considered the problem of the shock wave (SW) stability in media with arbitrary thermodynamic properties [1]. He developed the linear theory of the plane SW stability on a basis of the normalmode analysis and obtained simple quantitative criteria for different types of the shock behavior: -1 < 

L

 < 

L

0

(absolute stability),

L

0

 < 

L

 < 1+2

M

(neutral stability),

L

 < -1 or

L

 > 1+2

M

(instability). The notations are the same as in [1], correct expression for the lower boundary of the neutral stability region L0 was given by V. Kontorovich in [2]. The same results were obtained in works [3-5] differing from [1] in the features of the linearized formulation and/or used mathematical technique. More recently (see, e.g., review [6]), it was found that the Hugoniot fragments meeting to the instability conditions lie within the region of the SW ambiguous representation. The solution choice in such regions is of great interest, the more so the feasibility of the unstable SW with an unlimited growth of the amplitude of the front perturbations raises justified doubts [6]. It is also problematic the spontaneous sound emission by the front of the neutrally stable SW predicted by the linear theory. The attempts of the theoretical analysis of the neutrally stable SW behavior in certain nonlinear formulations [6-7] gave sufficiently presumable and discrepant results. Besides, the neutral stable or unstable SW (at least, in the form predicted by the linear theory) are not yet observed in experiments. It is clear that the problem considered goes beyond the linear theory and must be solved in the framework of as far as possible complete nonlinear formulation. In the paper presented we briefly review results of our systematic nonlinear analysis of the plane SW stability problem. Due to space restrictions some details of problem formulations and solutions we are obliged to omit, but they may be found in our works [8-10] The emphasis is placed on the origin of the cellular shock front structure arising if the SW has ambiguous representation caused by the fulfilment of the instability condition

L

 > 1+2

M

(this phenomenon has been first found in [10]).

It is well known that a shock wave can be attenuated when propagating through a porous barrier. Better understanding of this phenomenon is of significant importance in practical applications, e. g. in problems of industrial safety. As a result, the interaction of shock waves with a porous barrier has been studied in many papers. In particular, in [1, 2] a theoretical model, which assumes that a discontinuity breakdown happens at the moment when the propagating shock wave hits the barrier, has been developed. In accordance with this model the flow at large times is self-similar and the intensities of reflected and transmitted shock waves can be determined provided that some model is adopted for the flow through the barrier itself (typically it is considered as a steady subsonic flow).

Guderley theoretical model [1, 2] for weak shock wave reflection is a well-known way to overcome the von Neumann paradox within the gas dynamic framework. Recent Euler calculations [3] confirmed conceptual issues of Guderley theory for steady shock reflection. An expansion fan and a local supersonic patch were found behind the triple point. A numerical simulation based on the shock fitting technique was used in that study. Ivanov et al. [4], however, did not reveal any supersonic patches in Euler computations based on both the shock-fitting and shock-capturing techniques. Tesdall et al. [5] discovered a sequence of supersonic patches and triple points along the Mach stem for unsteady shock reflection. The results of numerical simulation showed a sequence of triple points and tiny supersonic patches behind the leading triple point. Another interesting result was also obtained in [6] within the framework of depth-averaged two-dimensional inviscid shallow water flow model. The nested-block grid refinement technique was used in those studies to achieve high resolution of the computational mesh. A supercritical patch was discovered. Thus, at the present time, the question about the supersonic patch structure is still open.

All previous studies on shock wave diffraction in shock tubes have spatial and temporal limitations due to the size of the test sections. These limitations result from either the reflection of the expansion wave, generated at the corner, from the top wall and/or of the reflection of the incident diffracted shock from the bottom wall of the test section passing back through the region of interest. This has limited the study of the evolution of the shear layer and its associated vortex which forms a relatively small region of the flow behind the shock and yet is a region of significant interest. A special shock tube is used in the current tests which allows evolution of the flow to be examined at a scale about an order of magnitude larger than in previously published results, with shear layer lengths of up to 250 mm being achieved. Tests were conducted at Mach numbers from 1.4 to 1.6 with wall angles of 10, 20, 30 and 90°

A mismatch has been shown to exist between the shock wave reflection behaviour on a circular arc and that on a plane wall at the same angle of incidence [1]. For reflection off a plane wall, this change in reflection pattern from regular to Mach reflection occurs at a wall angle where the flow behind the reflection point is just sonic in a frame of reference fixed in the reflection point. This transition condition is labeled the sonic criterion [2]. Experiments confirm this (except for some minor effects due to boundary layer growth). Data for convex cylindrical surfaces has shown that the visible eruption of a Mach stem, which is taken to be evidence of the point of transition, occurs at smaller wall angles than for plane walls [3]. On the other hand, recent tests [1] using perturbations generated by shock passage over a very small step in the wall have shown that the sonic catchup occurs at wall angles not only larger than those found for the visible eruption but also larger than the angles associated with sonic catch-up for the plane wall case. Due to the relatively low resolution of the time-resolved imaging system used, the accuracy of identifying the angle of sonic catchup could be questioned in the same way as identifying the visible eruption of the Mach stem could be resolution dependent. The stem could be erupting earlier if beyond the imaging resolution and the catch up could be occurring later bringing them both closer to the plane wall case. Due to this limitation time resolved tests were then conducted on a 45. plane wall preceded by a 75mm circular arc and the change in reflection angle measured in order to confirm that the reflection was different for the same incidence angle on the two different surfaces [1], this technique not being dependent on using perturbations. The existence of a significant transition length was confirmed in the reflection pattern adjusting from the curved wall to an eventual pseudo-stationary pattern on the plane wall. The results were still limited due to imaging resolution issues so a new set of tests have been undertaken using high resolution single shot tests over a range of wall angles and a range of combined surfaces. An additional test was conducted with a circular surface on both ends of a plane wall to explore how information from each of the joints influence the reflected shock.

A high-speed train entering a tunnel generates a compression wave that propagates through toward its exit.When the compression wave reaches the tunnel exit, a pressure pulse (“micro-pressure wave” [1, 2]) is radiated from the exit portal, and it causes an environmental problem. The magnitude of the micro-pressure wave is approximately proportional to the maximum pressure gradient ∂ 

p

/ ∂ 

t

max

(

p

: acoustic pressure,

t

: time) of the compression wave arriving at the tunnel exit [1].

Shock waves are formed when there is a sudden release of energy in limited space resulting in a supersonic displacement of the gas. Shock tubes are one of the easiest ways to generate good repeatable shock waves in ground-based test facilities with a good control over its parameters. A classical shock tube consists of a driver section(filled with high pressure gas) and a driven section(filled with low pressure gas) separated by a metal diaphragm. The rupture of the metal diaphragm generates a shock wave in the driven section of the shock tube. This mechanism of shock generation has been used for many decades and has been well investigated[1]. For classical aerodynamics studies, shock tubes of circular, rectangular or square crosssection and length scales of the order 50-100 mm are usually used. But the applications of shock wave assisted techniques in new areas like industry and medicine has led to reduction in the diameter of shock tubes and requires the understanding of shock tube flow at different length scales. The Ideal shock tube theory is based on assumptions like inviscid adiabatic flow, instantaneous diaphragm rupture and ignores aspects like diameter of the tube, surface roughness, boundary layer effects, heat and mass transfer effects, non-uniformity in driver section (Eg: combustion driven shock tube), possibilities of combustion at the contact surface and chemical kinetic effects. These effects which are ignored at macro-scales play a very important role when the diameter of the shock tubes are reduced and give rise to very interesting gasdynamic phenomena.

The study of thermo-chemical relaxation processes in shock tubes requires flow with sufficient residence time to allow the relaxation process to complete. In the context of hypersonic reentry, the length scales of interest are determined by the size of the flight vehicles, which in general are too large to be reproduced in the laboratory. For non-equilibrium binary kinetic processes, density-length scaling (where the product of density with a characteristic length scale is conserved) may be used to reproduce flight conditions. However, for situations where equilibrium is reached, any change to the overall pressure level will change the chemical composition, and similarity with flight will not occur, Morgan [1]. Therefore to study the radiation from regions of equilibriumflow, test gas at the same pressure as exists in flight must be used. Because the conditions at equilibrium are not path dependent, provided the appropriate pressure and temperature are reached, it is possible to study the radiation from an equilibrium region by creating just a small section of the flow field.

When a planar shock wave propagating down a channel encounters a decrease in cross-sectional area, not only is the shock strengthened but the shock shape and post-shock flow are disturbed. The current research investigates how various area reduction profiles affect the shock strength and shape, as well as the uniformity of the post-shock flow for planar shocks. Bird [1] and Russell [2] investigated experimentally the effects of wall shaping and strengthening by convergence respectively using strong incident shock waves; however the current study looks at relatively low Mach numbers of 1 < 

M

 < 2. By optimising the wall shape an area reduction can be used to increase the strength of a shock significantly without compromising on the quality of the post-shock flow, which is of particular importance in the design of experimental shock tube testing. More importantly, by analysing different profiles using numerical studies, the technique could then potentially be generalized to examine to what extent a shock shape may be purposefully manipulated to a required profile by suitable wall shaping. In order to provide a comprehensive study of the topic, numerical, analytical and experimental analyses are conducted. A comparison of computational fluid dynamic (CFD) simulations and Milton’s corrected ray-shock theory [3] is examined in detail.

The three-shock configuration is quite typical for the aircraft internal and external aerodynamics. For example, it appears at supersonic steady free-stream conditions in intakes (Fig. 1,

left

) and unsteady shock wave reflection from 2D wedges (Fig. 1,

right

). In the latter case the resulting flow is experimentally proven to be pseudo-stationary, and the triple point trajectory makes the constant angle

χ

with respect to the 2D wedge surface.

Attached-to-detached-shock transition analysis is of particular interest because the transition introduces radical changes in the flowfield and abrupt changes in flow characteristics. Section 2 of this paper discusses the possible causes of shock detachment from a sharp curved wedge. The concepts of

local and global choking

are introduced as causes for shock detachment from a sharp curved leading edge. Then the analytical conditions are shown under which the above causes are produced. Section 3 is a CFD confirmation showing flow field computations of attachment/detachment with the unstructured adaptive finite-volume Euler code SolverII [4] at the analytically predicted conditions. Two items are introduced here to aid the subsequent development of discussion, theory and computation.

Starting from the mid-1980s and up to now, experimental studies of the extracorporeal shock wave lithotripsy (ESWL) applications are performed in different fields of medical sciences, such as the treatment of kidney stone disease, neurosurgery, assisted drug delivery, the treatment of cerebral embolism, orthopedics and in the veterinary medicine. A generator of focused shock waves based on a high current spark discharge in water was developed in the Institute of Plasma Physics AS CR (IPP) [1, 2]. Hospitals in the Czech and Slovak Republic are equipped by such generators produced by the company MEDIPO, Brno. Number of patients undergo a course of medical treatment by these devices every year.

In hypersonic flow over a sphere, the shock wave stand–off distance is related to the density ratio by

$$\Delta/r\,=\,0.78\,\rho_\infty/\overline\rho,$$

where Δ is the stand–off distance,

r

is the sphere radius,

ρ

 ∞ 

is the free–stream density and

$\overline\rho$

is the average density along the stagnation streamline, see

e. g.

, Hornung [1]. For a sharp cone of a given angle, the stand–off distance increases linearly with

$\rho_\infty/\overline\rho$

from a critical onset point and is scaled by the base radius

R

of the cone, see

e. g.

, Leyva [2]. For a spherically blunted cone, one may therefore expect a transition to occur between sphere behavior and sharp cone behavior. This would be undesirable for stability and heat–flux reasons. A possibly more benign shape is an oblate ellipsoid, as suggested by Brown [5]. We study features of such flows on the basis of perfect–gas computations. Since the density ratio is very sensitive to reaction rate in flows with vibrational and chemical relaxation, these phenomena are very important in entry of vehicles into atmospheres such as that of Mars. The computations are therefore extended to include the effects of reacting CO

2

flows.

Within the joint project PAK 75 (Deutsche Forschungsgesellschaft DFG), a novel process for the production of gas phase synthesized particles is developed. The key features of the process are shock induced precursor combustion, homogeneous particle growth at constant thermodynamic conditions and gas dynamic nozzle quenching within a double choked Laval nozzle system (fig. 1) [1] [2].

In the design of supersonic airplane and air intake shapes, for specific performance, it is useful to begin with a known shock wave shape and flow-field and deduce the required wall shapes - design methods referred to as “wave rider” or “wave trapper”. Questions then arise as to the nature and existence of flow behind a given shock shape. The left lobe of a hyperbola of revolution shape is proposed as a particular example of a doubly curved, concave axisymmetric shock surface. It offers an analytically simple surface for the study of pressure gradient and flow curvature effects on shock detachment and reflection where the cumulative effects of both shock curvatures are present. Such shock shapes are physically plausible for internal, converging flow and Mach disk shapes. In this paper the concave, hyperbolically shaped shock in both planar and axial flow is investigated analytically with oblique shock theory as well as curved shock theory to discover any tendency towards the formation of a shock wave in the flow immediately behind the hyperbolic shock. If such a shock appears, and impinges on the back of the shock, then there would have to be a kink in the originally posed smooth shock and a Mach interaction would ensue. The onset of Mach interaction, at the sonic point is shown to depend on the free stream Mach number and the ratio of shock curvatures. Critical roles are attributed to both the subsonic patch of flow behind the strong portion of the shock and the orientation of the sonic surface at the shock. There is much experimental evidence of the existence of strong concave shock waves in the studies of Mach reflection where such shocks constitute the Mach stem. No experimental or CFD examples of continuously curved concave shocks that span both the weak and strong shock range have been found; probably because the enclosing ducts have to have very special shapes. Such surface shapes (both planar and axial) are suggested here.

Reflected shock interaction with an incoming boundary layer produces a complex, unsteady, three-dimensional flow field. Shock bifurcation, formation of recirculation bubbles, and turbulent jets are all observed and have been extensively studied experimentally ([1, 2, 3, 4, 5, 6]). The details of the reflection are known to depend on the inflow conditions, including the boundary layer behind the incident shock, and the wall boundary conditions. Reflected shock tube experiments have been conducted in shock tubes with both circular and rectangular cross-sections. There is experimental and numerical evidence that the bifurcated structure is substantially more complex near the corners of a rectangular tube as compared to the bifurcated structure on the centerline of a rectangular tube or in a round tube ([7, 8]). In this study, we present and analyze results of three-dimensional Navier-Stokes direct numerical simulations (DNS) of shock reflection in a square channel for three different incident shockMach numbers. Key features of the present simulations are very high resolution inside the boundary layer and temperature-dependent material and transport properties. We compare and contrast our results as a function of the incident shock Mach number with the existing theoretical model of Mark [1]. The simulations reveal additional flow features in the recirculation and corner regions that are not captured by the model.

Shock wave interaction with obstacles of various geometric shapes has always attracted attention in a large number of experimental and numerical studies. During the interaction of a shock wave with an obstacle a very complex wave pattern is formed which affects the shock-wave induced flow. The interaction reduces the shock-wave strength and generates rotational flow behind the obstacle. The interaction of shock waves with rigid obstacles is of significant importance in aerodynamic science and other engineering applications. Whitham [1] formulated an approximate theory for the dynamics of two- and three-dimensional shock waves and applied this theory to the description of shock diffraction by wedges and corners. Bryson & Gross [2] broadened Whitham’s theory and applied it to two- and three-dimensional bodies such as cylinders and spheres. They carried out theoretical and experimental work to assess the analytical computations that were made by Whitham. One dominant direction in investigation of shock-cylinder interaction is finding the RR→MR transition criterion. When the shock wave strikes a cylinder, it is reflected as an RR and then transforms to a Mach reflection MR. Major RR→MR transition criteria were summarized and discussed in a scientific monograph by Ben-Dor [3]. Since 1970, due to progress in numerical techniques, very accurate simulations of shock wave propagation over obstacles have been achieved. In most of studies efforts to validate the Euler scheme were undertaken. In the numerical study of Drikakis et al. [4] viscous effects were examined at various Mach numbers during of shock-cylinder interaction by comparing the inviscid and viscous calculations. It was found that the flow field in the downstream half of the cylinder is influenced by viscosity. The main objective of the present study is to better understand the physical elements governing the flow induced by the shock wave and the elements affecting the shock wave strength after passing the obstacle. To carry out the overall research plan two different approaches have been utilized - experimental and numerical. In the present study we focused on the investigation of the reflected shock wave from a single cylinder for low Mach numbers (

M

S

~1 − 1.4) in order to characterize the physical factors affecting its propagation. The first part of a broad investigation of the shock wave interaction with complex geometries is presented.

The interaction of shock waves with rigid obstacles is of significant interest in aerodynamic science and other engineering applications. During the interaction of a shock wave with an obstacle, a very complex wave pattern which affects the shockwave induced flow is formed. The interaction process depends on a variety of physical parameters such as the shape of the obstacle, the shock wave strength and the type of gas in which the interaction occurs. In the present paper, the interaction of a planar shock wave with a cylinder and a sphere is investigated. Our investigation follows closely the recentwork of Sadot

et al

. [1] which dealt with shock tube experiments with low Mach number shocks, in the range 1.1 to 1.4. An empirical relation was proposed for the trajectory of the reflected wave. This relation was expressed in terms of non-dimensional distance and time and was shown to be applicable for the investigated range of Mach numbers, cylinder diameters and a general ideal gas. The purpose of the present work is to focus on the backward reflected wave, and in particular, on its velocity change as it progresses away from the leading edge of the cylinder/sphere. It is expected that the reflected wave initially propagates at the velocity of shock reflection from a rigid wall, and asymptotically decelerates to the velocity corresponding to that of a sonic wave in the shocked region. This theoretical behavior is born out by fine mesh hydro-code computations of the interaction problem. The paper is organized as follows: in Section 2 a theoretical background for the limiting velocities of the reflected shock is given, followed by a brief description of the numerical codes and the problem setup (Section 3). The results of simulations for various cases by different CFD codes are given in Section 4. We conclude (Section 5) with a summary and suggestion of future work.

When a moving shock wave encounters a convex cylinder, reflects from it regularly, and propagates further, at one particular shock position corresponding to the so-called

sonic point

the flow on the cylinder’s surface, just behind the reflected shock becomes sonic with respect to the moving reflection point. The sonic point is prominent in the theory of regular-to-Mach reflection transition as one of its possible criteria [1]. When the flow behind the reflected shock wave becomes sonic, downstream perturbations can reach the reflection point and, supposedly,may cause the regular-to-Mach reflection transition.

We consider here the problem of a plane shock wave propagating in a turbulent flow field for the general purpose of investigating phenomenon of shock-turbulence interaction [1]. We compute this using a molecular kinetic model theory approach with the Boltzmann equation and the BGK approximation [2]. In practice, we first prepare an advancing plane shock wave in uniform gas made by shock-tube-flow computation as shown in the pressure distribution in Fig. 1 (right) and then run this into a prefabricated isotropic turbulent field in a square region from one end (Fig. 2). Particular attention is paid to the increasing shock front thickness compared to the shock in a non-turbulent field (see Fig. 3 ). This thickness is compared with the shock thickness given by mixing-length theory [3] and observational data of a shock-tube-wind-tunnel experiment [4].

A supersonic jet is known as a source of remarkable noise which can become intolerable in the neighbourhood of jet engines, cutting torches and other supersonic jet producing devices. The most dangerous part of the noise is emitted in form of Mach waves. They appear inside as well as outside of the jet and are quite regular. The practically ideally expanded supersonic free jet in Figure 1 left hand side shows outside Mach waves visualized by Oertel sen. [1] using a differential interferometer. The Mach waves inside can best be detected by immobilisation records on moving film as shown in Figure 1 on the right hand photo with the w’-Mach waves outside, the w-Mach waves inside and outside and the w”-Mach waves inside.

Interaction between a shock wave and turbulence gives challenging research problems. Interactions between a shock wave and a vortex or a flow element of the same kind have been studied in various manner.[1-8] Impacts of a shock wave to turbulence intensity is investigated by Honkan et al.[1] and Barre et al.[2]. On the other hand, the impact of isotropic turbulence on a shock wave has been studied mainly numerically[8]. Experimental data of shock wave and isotropic turbulence are currently insufficient. The purpose of this study is to obtain experimental data on the relationship between a post-shock overpressure and characteristics of isotropic turbulence.

Parachutes are often used for planetary-entry and re-entry. Most of cases, parachutes are deployed at supersonic speed. Therefore, it is necessary to obtain aerodynamic characteristics at supersonic speed. However parachutes have wrinkles and seams. They make it difficult to study aerodynamic characteristics of the parachute shape. Therefore, experimental studies using solid models have been conducted from the 1960s re-entry program of the US [1]. In particular, Helmut G. Heinrich et al. visualized shockwaves, and shockwave vibrations in Mach number 3 [2]. These shockwave vibrations make the flow unstable, and the unstable flow gives model vibrations. They are risks of break.Therefore,Helmut G. Heinrich suggested an improved model.

Vortex interaction with a rotating blade is fundamentally important in understanding the unsteady aerodynamics and aeroacoustics problems arising in a helicopter flight. The blade-vortex interaction (BVI) can occur when the main- and/or tail-rotor blades interact with tip vortices previously shed by preceding high speed blades of the main rotor. Most numerical studies have focused on limiting cases where an idealised vortex interacts with a single, non-rotating blade. Despite the significant differences between a real case scenario of BVIs involving more complex characteristics of blade tip vortices and rotating blades, numerical and computational studies nevertheless provide great insight into the deeper understanding of the fluid mechanics that are relevant to the problem. Our computational modelling approaches this problem in several steps. Firstly the ‘impulsive’ instantaneous blocking of the column vortex by a flat plate. This was studied theoretically for incompressible motion byMarshall [1], who produced an elegant model for ‘area change’ waves in the vortex core. Our Computational Fluid Dynamics (CFD) study essentially models this, together with the true compressible pressure wave formation that also arises. In the second step we extend this to the gradual cutting of the vortex by a sharp flat plate that moves, at a finite speed, through the vortex. Finally, we extend this to the gradual cutting of the vortex by a blunt leading edge aerofoil that moves, at a finite speed, through the vortex, which incorporates both the ‘blocking’ and also stretching and distortion of the vortex lines. In this paper, we present only the 3D numerical computations of a single columnar vortex instantaneously ‘blocked’ by a flat plate.

The family of the difference schemes on a minimal stencil is under consideration. Construction of the difference schemes on a minimal stencil is based on the scheme approximation order increasing procedure [1]. This method makes it principally possible to develop the schemes of arbitrary approximation order without extension of the scheme stencil by the use of differential consequences of the initial system of equations. Two-dimensional schemes of similar type for plane and cylinder flow symmetry were presented in [2] - [4]. The schemes on the flow oriented grids for plane, cylinder and spherical flow symmetry supplemented by shock-tracking procedures were presented in [5]. Validation and comparison of calculations with the use of the minimal stencil difference schemes and the other ones was conducted in [5] - [7].

In hypersonic flow over slender bodies an important concern is the influence of nose bluntness on the viscous boundary layer behavior. This is particularly important in the consideration of boundary layer instability and transition. There exist numerous treatments, both theoretical and experimental on this subject. Representative excellent examples are [1] and [2] In this study we use the deviation from sharp–body theory of the heat flux distribution as obtained from computations of viscous perfect–gas flow over blunted slender bodies at zero incidence to estimate the point where the body has forgotten that it is blunt.

Shock/Vortex Interaction is one of the fundamental problems of aerogasdynamics, which has been adequately studied neither theoretically nor experimentally [1, 2].

One of the main features of the vortex-shock interaction is its severe unsteadiness. It is experimentally observed in fluctuations of gas-dynamic parameters of the flow and changes in the flow structure and the size of the interaction region. Experimental data with quantitative estimates of manifestations of unsteadiness are rather limited. The absence of numerical and experimental data for hypersonic velocities should be specially noted. Obtaining results for this range of velocities is extremely important for the development of promising flying vehicles (avoiding of catastrophic operation regimes of a hypersonic inlet and improvement of mixing in the combustor).

The flow field dynamics associated with blast waves can be better understood by generating controlled micro-explosions in the laboratory. In recent years microexplosions have also found interesting trans-disciplinary applications like food preservation, wood science, drug delivery, gene therapy and bio-medical applications [1], [2]. The blast waves produced by sudden release of energy are normally characterized by a supersonic shock front followed by an exponential type decay of its physical properties. Unlike shock waves that attenuate as they expand spherically, the shock wave from an internal blast can change its propagation properties depending on the physical barriers. The micro-blast provides a challenging case for application of novel flow diagnostic techniques in measuring flow properties. This learning can be scaled up to large scale explosions [3]. One such property is the density field, which although highly informative, is quite difficult to capture. The Background Oriented Schlieren (BOS) technique provides the capability of capturing the three dimensional density fields [4], [5]. This is an attempt to quantify the density flow field of a micro-explosion for the first time using BOS. In this study, a micro-explosion is generated using NONEL tube and the detonating device. The spatio-temporally evolving density field is captured at several instants by means of a precise triggering circuit used to control the illumination and imaging. The density field so obtained can be used for understanding both basic physics associated with explosive driven shock wave propagation as well as validation data attempts to model explosive driven shock wave propagation.

The environmental destruction of marine ecosystem caused by micro-organism included in ship ballast water has been a global problem. The International Maritime Organization (IMO) adopted strict standard rule for control and management of ship ballast water in 2004 [1], after that, many ballast water treatment systems have been proposed and developed all over the world. However, a lot of practical problems have been remaining yet for development of energy-saving and space-saving systems. In general systems, shipping companies have to charge and manage chemicals on board to kill marine bacteria in the ballast water. Therefore, if troubles should happen in the ballast water treatment process, leaking chemicals from the system might contaminate the sea. In order to realize more secure and environmental friendly treatment method, the authors have proposed a new sterilization technique of ship ballast water using underwater shock waves. In the previous research, the result showed that a marine

Vibrio

sp. was completely inactivated when the excess pressure in cell solution contained in a small aluminum container was over 200 MPa in impact experiments by a gas gun [2]. In general, electric discharge, explosive or high-speed collision of a projectile is used as a power source to produce underwater shock waves and has been applied to engineering and medical fields. However, those power sources are unsuitable for practical use on board from a point of view on energy cost and safety. From the above-mentioned, the authors thought of killing marine bacteria in a large amount of ballast water by exposing to strong pressure pulses and free radicals created from collapse of microbubbles [3-6]. In this idea, the excess pressure of underwater shock waves plays only a role of leading to collapse of microbubbles, so that it does not necessarily need to release an extreme high-energy in water. Therefore, the method of hitting water surface with shock waves produced in gas would be an effective one of underwater shock wave generation on board. In addition, it needs to develop the shock wave generator that can produce shock waves periodically by low driving cost for practical use.

When a shock wave propagates over many small solid particles on a horizontal wall, some particles near the surface of the layer are lifted and dispersed into the shockinduced flow. These dispersed particles is called the dust cloud. This phenomenon is actually seen in galleries of coal mines or in pipelines for neumatic transportation of powder, and mixing dispersed flammable dust particles with high-temprature and high-pressure gas behind the shock wave sometimes causes the dust explosion. And this phenomenon includes some interesting factors, such as the shock structures interacted with the dust layer, interactions between gas and solid particle, and interactions between solid particles.

Contaminant metal particles of the order of 100-500 microns in diameter in the liquid propellant feed systems of rocket engines are a significant hazard and safety concern. These particles may originate from within the propellant tanks, valves, feed lines, pumps, or the propellant itself. Ignition and combustion of the particles when located within a supercritical oxygen-rich environment, such as would be found in an oxidizer-rich rocket engine system, could release a significant amount of energy. In addition, particle impacts with the walls of the propellant feed systems could sufficiently heat the particles to ignite them or to fracture them into smaller particles that are easier to subsequently ignite. Experiments are currently underway at The Aerospace Corporation to improve the knowledge base for particle-impact ignition. Oxygen pressure, particle size and kinetic energy, and the occurrence of fragmentation upon impact are among the parameters to be studied where impacting particle velocity is induced by the passing of a shock wave [1]. A simulation capability, validated by the experiments, would be of value to predict the risk of possible particle contamination.

The starting process of supersonic planar nozzles has been the subject of great number of the shock tube researches in the past. Initially this was motivated by the need to clearly separate the unsteady and quasi-steady parts of the expansion flow and thus specify the so-called ”test time” period of shock tube tunnels. Among other, the best known images illustrating the starting process were published by Amann [1]. It was clearly shown that the starting flow initiated by the primary shock wave (PS) includes the contact surface (CS) and the secondary shock (SS). Smith [2] was the first to show that the unsteady expansion wave (UEW) which follows the SS can also affect the total duration of the starting flow. Actually, the SS initiates flow separation and transient structure of the separation points (SP). Next, complex phenomenon which requires fundamental knowledge on the parameters of the external flow and the condition inside the boundary layer was discussed by Dussauge & Piponniau [3]. Flow separation may also cause significant effects on the trajectory of the SS and increase the total duration of the starting flow pattern. The renewed interest in nozzle starting phenomena appears due to wide application of the transient nozzle flow in different devices. The effect of separation, for example, becomes important inside the nozzles of rockets, missiles and/or supersonic aircrafts where it is usually undesirable since it may cause a dangerous lateral force which can damage the nozzle [4]. On the other hand, flow separation and the resulting instability of the exit plume could have positive effect when used in high speed mixing devices (Jonson & Papamoschou [5]). Despite a plentiful history and significant progress in the numerical as well as in the experimental investigations, many features of the nozzle starting, flow separation and its asymmetry are still open. Even a brief summary of the involved process clearly shows that the nozzle geometry, the viscous effects and the flow conditions at the entrance and exit are important parameters that must be examined [6]. The current investigation was conducted to evaluate how the nozzle starting process depends on the initial conditions and on the asymmetry of the nozzle installation. In order to assess the role of these factors the incident shock wave Mach number was varied between

M

s

 = 1.2 and

M

s

 = 1.9 and the nozzle asymmetry was introduced by the relative shifting of half of the nozzle.

Since the pioneering work by Ziemer et al.[1], the interaction of the weakly ionized flow with a magnetized body has been investigated both experimentally[3, 4, 5, 6, 7, 2, 8] and numerically[9, 10, 11, 12]. This interaction has become a topic of interest as it has attractive applications to the mitigation of aerodynamic heating in hypersonic flight vehicles, which was numerically demonstrated by Poggie[9] and known as the electrodynamic heat-shield. Thus far, except for limited reports[13, 8], mitigation of aerodynamic heating has not been experimentally demonstrated. One such recent report was made by Gülhan, but its result is still debatable[14]. Alongside the experimental efforts, numerical investigations have been conducted intensively[10, 11, 12]. It has come to our attention that, unlike the simple flow model assumed by Poggie[9], the interaction may be influenced by a variety of effects such as the Hall effect [10].

At present, there are several groups of spacecraft (THEMIS, Cluster, Double Star) rotating around the Earth on orbits whose elements are located in the outer magnetosphere, magnetosheath, and in the neighborhood of the Earth’s bow shock

S

b

at distances of 10÷25 Earth’s radii

R

E

. Other spacecraft (Wind, SOHO, ACE) are located in the free solar wind in the neighborhood of the Lagrange point

L

1

at ~250

R

E

from the Earth. The data on the solar wind, the interplanetary magnetic field (IMF) and the magnetosheath parameters are continuously transferred to the Earth and actively analyzed. The measurement results are used to identify sharp changes in the solar wind associated with shock waves and other discontinuities and their manifestations in near space [1–4]. These investigations are due to the need for forecasting the cosmic weather which manifests itself on the Earth in the form of sudden storm commencements, magnetic substorms, and sudden impulses.

The present work is concerned with the differences in how shock and detonation waves inside pipes or ducts reflect from closed ends. One of the motivations for the present study is that the large pressure rise associated with a detonation poses a hazard to pipes that contain flammable mixtures [1]. A detonation impinging normally on a planar wall creates a reflected shockwave to bring the flowat the wall to rest [2] and produces pressures 2.4 times that of an incident Chapman-Jouguet (CJ) detonation [3]. In examining the material deformation produced by reflected detonation loading [4] an inconsistency was discovered between the measured pressure jump across the reflected shock wave and the measured speed of the shock, with the measured pressure being as much as 25% below that predicted by the shock jump relations for the given shock speed. This was theorized to be due to bifurcation of the reflected shock wave associated with shock-wave boundary layer interaction.

The study of cylindrical and spherical converging shock waves propagating in solid materials is relevant to the production of high temperatures and pressures in condensed matter with applications to inertial confinement fusion [1]. However, experimental studies conducted in the area are prone to complications derived from the measurement techniques available and the difficulty of producing a quasi-radially symmetric flow.

The thermoelectric material is a material that shows large thermo power, low resistivity and low thermal conductivity. Recently, layered cobalt oxides have been extensively investigated as a promising candidate as thermoelectric material. Important feature of sodium cobalt oxides is that, the sodium ions (

Na

 + 

) randomly occupy the regular site by 50% and the sodium content can go up to 70% changes[1]. In this sense layered cobalt oxides, Na

Co

2

O

4

should be written as

Na

x

CoO

2

(x = 0.5) and the compound is quite promising for thermoelectric power generation. Terasaki et al. found that a single crystal

NaCo

2

O

4

exhibits good thermoelectric property[2]. These layered oxides consist of two layers:

CoO

2

layer and Na ion layer.

CoO

2

layers acting as an electron reservoir which are responsible for the electrical conductivity and large thermoelectric power. Na ions layer sandwiched between two neighboring

CoO

2

layers adjust the concentration of electron in

CoO

2

layers and decrease the thermal conductivity along the stacking direction c[3]. Since the discovery of moderately large thermoelectric power (Seebeck coefficient) together with high electrical conductivity in

Na

x

CoO

2

(x is 0.70), experiments were done to find new phases for thermoelectric conversion applications[4]. A single-crystal X-ray diffraction study confirmed that

Na

x

CoO

2

(x = 0.74) adopted the hexagonal P6

3

/mmc space group[5]. Neutron diffraction and electron diffraction study shows that the crystal structure of the oxides is strongly dependent on sodium content. The crystal structure of

Na

0.75

CoO

2

was studied at ambient and low temperatures down to 10 K at pressures up to 40 GPa using a diamond cell shows an increase in Co-O bond length and decrease in Na-O bond length[6]. Here we present the experimental results on the interaction of shock heated test gases with

Na

x

CoO

2

at high temperature and moderate reflected shock pressure.

The behaviour of shock waves propagating in a gas has been studied for more than a century, experimentally, theoretically and numerically. It was established, in particular, that the thickness of a shock wave (the layer, in which parameters of the medium vary rapidly in space) is of the order of several mean free paths of the gas molecules, provided that the shock is of moderate intensity. Distributions of the parameters of the medium inside the shock resemble the hyperbolic tangent function. The necessary condition for the above to be fulfilled is that the gas is dilute, which means that its molecules interact (collide) with only one neighbour at a time, and between collisions they move with constant speed along straight lines (concept of a mean free path – the average distance travelled this way). Duration of a single collision is negligibly short as compared to the time of free flight (mean free time). In dilute gas the mean free path of the molecules is much larger than the diameter of the molecule and larger than their average separation distance (Figure 1 – left).

Laser ablation is a frequently used method of removing material from a solid surface by irradiating it with a powerful laser beam. It may be applied to machining materials, cleaning contaminated surfaces, deposition of thin coatings on surfaces etc. High energy, short duration laser pulse, focused on a small area of the target surface heats and evaporates it, forming eventually a plume which moves outwards from the target with high speed. The behaviour of the plume may influence the quality of the deposited layer, which is important if deposition is the goal of the process. This is particularly the case if the deposited material consists of disparate mass components. The light components move faster than the heavy ones and tend to spread on larger area of the substrate. In consequence the stoichiometry of the deposited material is not preserved. To improve the situation, the deposition process may be performed in the atmosphere of an ambient gas, which decelerates both the motion of the plume as a whole and its expansion. Deceleration is stronger for light components of the plume, which makes the expanding plume more uniform.

Laser energy focused in bulk of transparent dielectric creates heated and elevated pressure zone producing nonreversible media destruction. Physical nature of this effect is steady in focus of modern physical studies since laser invention. Silica is quite attractive object for this due to its numerous physical and technical applications. Silica-based optical fibers provided new unique possibility for laser damaged zone study. Single mode optical fiber has constant laser radiation distribution in any transverse cross-section along the full length. This property makes available to observe propagation of energy deposition zone under steady-state conditions as opposed to laser focused into unknown initial volume of transparent dielectric target, particularly for long laser pulses (

τ

p

 > 1

ns

). The damage zone in optical fiber can expands to any distance to laser direction. Such laser driven wave (LDW) destroys the core and sometimes cladding too. Temperature in fiber optic core can achieve up to ~10

4

K

and silica goes to hot plasma with solid density.

At present, several groups of spacecraft in the vicinity of the Earth’s bow shock

S

b

, magnetosheath, and outer magnetosphere and spacecraft located in the free solar wind stream near the Lagrange point

L

1

are measuring the magnetosphere and solar wind parameters and the interplanetary magnetic field (IMF). The data obtained on the Earth are analyzed to forecast the cosmic weather, in particular, to identify sharp changes in the solar wind associated with shock waves [1, 2]. For interpreting adequately spacecraft’s measurements it is necessary to have exact solutions of the problem of interaction between a solar wind discontinuity and the bow shock in which the wave flow pattern is known as a function of coordinates of a point on

S

b

.

Laser energy deposition has brought great interest to researchers due to its applications in drag reduction [1, 2], shock wave modification [3, 4], fuel ignition [5], and optical perturber for transition study [6, 7, 8]. Compared to the electric discharge flow control techniques, laser energy deposition can excite the flow non-intrusively with almost any pulse-width and repetition rate [9] with out any electrodes.

In the oil industry, acoustic techniques are commonly practiced to determine the position and the properties of the reservoir and the overburden. These techniques comprise the use of seismic surveys, cross-well tomography, and borehole logging. In the latter technique, acoustic sources and detectors are installed in a logging tool that is run in the borehole penetrating the potential hydrocarbon reservoir (i.e., a porous rock formation). The acoustic source generates a variety of borehole wave modes among which the Stoneley wave is most prominent [1]. In the field, the borehole is usually intersected by natural reservoir fractures and faults that may extend over several kilometers and dramatically affect the borehole acoustics [2]. Here we use a conventional vertical shock tube to generate and study wave propagation in a borehole intersected by a single horizontal permeable fracture.

The research on shock-wave induced bio-effects is expanding rapidly where an emerging field is the so-called cell transformation, i.e., the uptake of deoxyribonucleic acid (DNA) from the surrounding. ‘Competent bacteria’ are those which are capable of being transformed. The standard method to identify transformed cells uses plasmids (DNA molecules that replicate independently of the chromosomal DNA) containing a gene that increases bacteria resistance to the antibiotic they are normally sensitive to. After plating the bacteria on a medium containing the antibiotic, only the transformed cells proliferate. Chilling the cells in CaCl

2

, shocking them with an electric field to create holes in the membrane, and ultrasound are physical methods to increase bacteria competence [1]. Nevertheless, there is still a lack of efficient methods for DNA delivery. Cell transformation by ultrasound is based on cavitation-induced membrane permeability [2]. Shock wave-induced sonoporation has also been associated with cavitation, i.e. growth and collapse of microbubbles [3-5]. In most studies clinical shock wave generators have been adapted to apply up to several hundredths of shock waves to a vial containing cells in suspension. After passage of each shock wave, a cloud of bubbles forms inside the vial. These bubbles expand and collapse violently after approximately

$250-500\ \mu$

s, emitting high speed microjets that are supposed to be responsible for cell transformation [6]. Microjet emission can be intensified if a second shock wave (Fig. 1) arrives shortly before the bubbles start to collapse. This phenomenon has been used to improve kidney stone fragmentation [7-8]. Temperature is another factor affecting cell transformation. Increased membrane permeability due to a temperature reduction has been reported [9]; however, if temperature reduction enhances microjet emission and thus contributes to cell transformation is unknown. To analyze this issue, the dynamics of a bubble immersed in water was simulated by using a well-known numerical model. The object of this study is to enhance shock wave-induced transfer of plasmids into

E. coli

using tandem shock waves and to analyze the influence of the temperature on bubble dynamics and cell transformation.

Drug Needle injectors have been the common means by which vaccines and protein therapeutics are transdermally delivered. However, the use of needle injectors have elicited painful reactions, and have also caused infection due to repeated use of needles particularly in under-developed countries [1]. Because of these disadvantages, researchers have endeavored to developing alternative methods for drug delivery. New methods incorporating liquid jet injections have been developed [2]; however, liquid jet has not succeeded in replacing needle based injectors yet. Despite the fact that jet injection can alleviate patients aversion to needles, it has not gained much popularity for the following reasons: i) it is still reported to be painful, ii) it is not stabilized in control, and iii) the risk of cross-contamination due to back splash is still prevalent [3, 4]. Therefore, in order to release the stronghold that needle injection has on the drug delivery domain, methods such as jet injection must make a major breakthrough.

As new biomedical and industrial applications of shock waves emerge, the need to accurately and economically generate shocks is becoming more critical. Since a very large potential resides in biology and medicine areas for diagnostic and therapeutic uses, shock waves need to be efficiently produced in cells, tissues and organs. In the past, there have been a number of methods used to produce shock waves in liquids, all characterized by a large and rapid energy deposition, either through the detonation of an explosive, the irradiation of a target with a pulse of laser energy, the dumping of electricity through a spark gap, or the sudden acceleration of a piston, either by electromagnetic or piezoelectric means. There are well known shortcomings associated with each of these methods, such as the requirement for high-voltage electronics, the manipulation of explosives and/or the lack of control over the shock properties [1]. This paper presents a new method to generate highamplitude pressure pulses in liquids exploiting the advantages of low amplitude piezoelectric generators.

Shock wave applications to medicine conducted in the Interdisciplinary ShockWave Laboratory of the Institute of Fluid Science, Tohoku University were based on accumulation of underwater shock wave research and bubble dynamic study, and initiated with the development of micro-explosion assisted Extracorporeal Shock Wave Lithotripsy (ESWL). The Ministry of Health Japan approved in 1987 our ESWL system as clinical device. The result was then continued to therapeutic devices in orthopedic surgery, gastro surgery, neurosurgery, drug delivery and recently cardio vascular treatments. In these projects, we have been working with colleagues in the School of Medicine, Tohoku University, and always target to develop therapeutic devices as a fruit of our shock wave dynamics and bubble dynamic studies.

Medical, industrial, and environmental applications of pulsed power technology have been developing rapidly in many fields including bioelectrics for cancer treatment and induction of apoptosis; treatment of exhaust gases; sterilization of microorganism; removal of biological wastes; fragmentation of rocks; recycling of concrete and electrical appliances; and surface treatments of material[1]. The application of electric fields with a short pulse width allows direct interaction with biological cells substructure without heating the tissue[2], which suggests interesting comparison and/or combination with shock waves for medical applications. The breakdown phenomena in liquids have been studied for a long time, in particular for its relation to electrical insulation.

In this presentation, development of medical and biological applications by shock waves and bubbles, especially drug delivery systems, regenerative therapy (angiogenesis) and water treatment, is explained. We have developed the drug delivery systems using shock waves from fundamental investigations. In this system, microcapsules including gas bubbles are flown in the blood vessel, and broken by shock induced microjet, and then drug is reached to the affected part in the body. For developing the microcapsules including gas bubbles, the penetration force of microjet should be controlled by shock wave, rise time of pressure history, and capsule geometry and material properties.

Extracorporeal shock wave lithotripsy (ESWL) has been successful for more than 30 years in non-invasive treating patients with stone deceases (mostly kidney stones). ESWL devices (lithotripters) generate shock waves outside the patient’s body and concentrate them on the kidney stone. Over 40 models of lithotripter (electrohydraulic, piezoelectric and electromagnetic) are commercially available worldwide [1],[2]. At the end of 1980’s one of modified versions of electrohydraulic type of lithotriptor was developed also in the Institute of Plasma Physics AS CR [3]-[5]. Such generators serve up to now as a therapeutic unit in the lithotripters Medilit (produced by the company MEDIPO, Brno, Czech Republic) at about 20 hospitals in the Czech and Slovak Republic [6]. So far more than 120 thousands patients have been successfully treated by these devices.

Repetitive-laser pulse energy depositions are contributed to reduce the wave drag of supersonic flight in this study. The intensive laser-heated gas generated by laser beam focusing is useful to control the supersonic flow field.When the laser energy is deposited into the air, blast waves and spherical laser heated gas interacts with bow shock wave in front of supersonic flight. Thereafter, laser-heated gas is transmitted to shock wave, and vortex is generated by baroclinic effects.

The idea to use the energy deposition, localized in supersonic flow upstream of a body, for the improvement of aerodynamic characteristics was proposed in Russia more than 20 years ago. Theoretically the effects of wave drag reduction and flow reorganization were observed for supersonic flow past sphere by Georgievskiy and Levin [1]. Experimentally the wave drag reduction of blunt and streamlined bodies was confirmed when the optical laser spark was realized in upstream supersonic flow in single pulse mode by Yuriev et al. [2] and in pulse-periodic quasi-stationary mode by Tretiyakov et al. [3]. The recent survey of flow control and aerodynamic drag reduction by the energy deposition was presented by Knight [4].

The performance of hypersonic propulsion can be critically affected by shock wave/boundary layer interactions (SBLIs), whose severe adverse pressure gradients can cause boundary layer separation. This phenomenon is very undesirable in engine intakes leading to total pressure loss and flow distortion which can cause engine unstart. Hence, it is essential to apply flow control method to the flow, either at the beginning or during the interaction phenomenon to prevent the shock-induced separation [1].

In the recent past, there is a growing interest in using the Dielectric Barrier Discharge (DBD) based plasma actuators for active control of the boundary layer in different speed regimes. The plasma actuators are attractive flow control devices because of their simplicity of construction and their very short response time. The plasma actuation relies on one or more of the three basic mechanisms viz., volumetric joule heating, electrohydrodynamic (EHD) forcing and magnetohydrodynamic (MHD) forcing. DBD actuation has been widely accepted as due to EHD force it generates. Beouf et.al[1] have discussed the basic mechanisms responsible for EHD force exerted by DBD. The EHD forces on the fluid generates a flow in the vicinity of the DBD. A planar configuration of DBD in flush with surface generates a wall jet as reported by Moreau[2]. The planar configuration has been extensively explored for aerodynamic flow control at subsonic speed.

A great number of researches in the recent years deal with the problem of non-equilibriums plasma flow control in aerodynamics [1]. Different approaches are used for efficient flow moderation with energy deposition in a boundary layer, compression wave area, duct inlets and others [1, 2, 3]. Energy input using pulse discharge plasmas appeared to be rather promising. Pulse and pulse-periodic energy supply is the most effective way to improve high-speed flow characteristics; shock waves arising from local pulse energy input area may influence high speed flow with shock configurations.

Historically the potential of energy-assisted shaping of high-speed flows with modest on board power requirements has been the subject of a number of earlier investigations. The possibility of obtaining drag reduction using energy sources upstream of blunt bodies has been pioneered by Georgievskii and Levin [1] and Myrabo and Raizer [2] in theoretical studies. In their studies the magnitude of the drag reduction was found to be insensitive to the location of energy deposition at a sufficiently large distance from the body. This was followed up by various computational studies; Levin and Terenteva [3] and Riggins et al. [4] showed power savings over cones and blunt bodies using two dimensional Euler/laminar computations.

It is known [1] that the interaction between shock waves and a discontinuity surface is poorly studied, in particular, in the presence of a local unsteady heat release in the flow. Increased interest in the propagation of shock waves in a moving medium is caused by the problems arising when studying supersonic burning in gas and optical discharges and the generation of control efforts on a flowed around surface with the organization of a heat release zone above it. An unsteady contact surface can exist, for example, in the wake of a pulse-periodic optical discharge. It is possible to single out the line recently developed and related to the possibility of attenuating a shock wave after the passage of the thermal layer formed by a pulse-periodic energy source [2, 3, 4, 5]. When the incident shock wave interacts with a contact discontinuity surface, transmitted and reflected disturbances arise.

Since the dawn of time, human kind have felt the presence of shock waves in nature through thunders and vulcano eruptions and, unable to understand them, have associated their often destructive might to divinities such as Zeus and Jupiter in the Greek and Roman mythology, the Norse divinity of Thor, and the elusive Thunderbird in the Native North American culture. Through history, albeit unknowingly, humans have been able to generate shock waves via the cracking of a whip or the explosion of fireworks. It was the invention of the atomic bomb that brought back fear and respect towards the might of this natural phenomena [1]. For research purposes, shock waves can be easily generated in a laboratory environment using shock-tubes where a high-to-low pressure discontinuity is initially present.

The drag and heat transfer are very important parameters to be taken into account in the design of hypersonic vehicles. Under certain conditions, spikes can substantially reduce the drag of axisymmetric blunt bodies at supersonic and hypersonic speeds with the deliberate use of flow separation. This happens mainly due to the induced formation of a conical shock in opposition to a normal one.

The authors have proposed a new method for ship ballast water treatment using shock pressure generated by collapse of microbubbles. Our previous study showed that the marine

Vibrio

sp. in an aluminum container was killed by exposing to the excess pressure higher than 200 MPa directly generated in the suspending medium by the impact of a projectile accelerated with high-pressure gas [1]. Applying the shock pressure sterilization to a huge volume of ballast water, the projectile impact method is not a practical way to generate shock waves successively in liquids. For the practical use, we have suggested the use of microbubbles for one of the alternative methods as a micro-generator of impulsive pressures. When microbubbles collapse, they generate shock pressures and free radicals that can oxidize something around them [2, 3]. However, it has been not well known about the inactivation effects of shock pressure produced from collapse of microbubbles on the marine bacteria. In order to find the effective conditions on the shock wave sterilization, it is important to investigate the motion of microbubbles exited by shock waves experimentally and analytically.

The trans-boundary movement of the microorganism included in the ship ballast water causes the destruction of an marine ecosystem globally. The restrictions concerning the ballast water managements provided by International Maritime Organization (IMO) probably take effect for all new ships in the near future, so that enterprises and research institutions in the world have developed many kinds of ballast water treatment systems to obtain the IMO approval. In most of all systems, chemical treatment is used to kill marine bacteria such as cholera and colon bacillus. For marine environment convention, safe and economic and eco-friendly ballast water treatment systems are desired.

A Richtmyer-Meshkov Instability (RMI) [1, 2] is generated when an interface between two different fluids is impulsively accelerated. The instability develops due to misalignment of the density and pressure interfaces. This misalignment results in the deposition of vorticity, causing the formation of an instability that grows nonlinearly with time and eventually may transition to fully turbulent flow. It has been recently shown that a similar class of instability can evolve in a multi-phase flow [3], where the density gradient is caused by a second, non-fluid phase.

The effects of viscosity and heat conduction, heat losses due to the wall heat transfer, as well as nonequilibrium phenomena can play an important role in microflows. Recent numerical investigations [1] of shock wave propagation in a microchannel with allowance for viscosity and rarefaction effects revealed significant differences from the inviscid theory, which ensures a correct description of the majority of specific features of macroflows. In that work, the shock wave was generated by breakdown of a diaphragm separating high-pressure and low-pressure domains.

For many applications where solid and heavy protections against blast are inoperative, the mitigation of the blast wave loading in a cost-effective manner could be achieved using aqueous foam. The protective behavior of aqueous foam is mainly ascribed to high compressibility of the gas bubbles, which is generally accomplished with energy losses due to side wall friction, viscous losses, evaporation, foam shattering and acceleration of the resulted droplets [1, 2, 3]. As transient processes, these factors introduce uncertainty into the predicted behavior of the foam based protection [4]. Recently it has been established that solid additives slow down the foam decay due to the increase in the liquid viscosity [5, 6] as well as enhance the mitigation performance of the foam barriers [7]. A diversity of physical mechanisms responsible for the final effect complicates the issue, and to obtain reliable data, one has to use specially designed tests.

Triple-shock-wave configurations (TC) were experimentally reported by E. Mach in 1878 [1]. The first detailed experimental and theoretical investigation of the TC was fulfilled by J. von Neumann in 1943 [2]. In the theoretical description of the TC in the von Neumann article [2] (known as the three-shock theory) gas was supposed thermally and calorically perfect. The three-shock theory is based on the Euler equations in the integral form, and expresses all non-dimensional TC parameters through the two basic ones: the initial flow Mach number

M

and either the incident shock intensity

J

1

or its inclination angle

σ

e

1

[3]. In [4] the problem of a moving shock wave reflection from a plane wedge with a TC formation in real gases (nitrogen, air and oxygen) was considered.

The Richtmyer-Meshkov (RM) instability [1, 2] occurs on a perturbed interface separating two fluids with different densities when impulsively accelerated, which has received much attention because of its academical significance in the field of vortex dynamics and turbulent mixing, and having important applications in inertial confinement fusion, supersonic combustion, and supernova collapse. In the past decades, scientists have performed many experimental, numerical and theoretical researches on RM instability. Numerically, the research work is usually carried out based on the experiment in order to further understand the instability phenomenon. Meyer and Blewett [3] employed a Lagrange algorithm to simulate the process of the RM instability. Later, the front tracking, high order WENO shock-capturing method and LES have also been developed to study the corresponding problems. However, the previous study mainly focused on the circle or single-mode interfaces. Recently, the shock interaction with a rectangular block was studied by Bates

et al

. [4]. Bai

et al

.

Following the original numerical work by Hornung, [1], an experimental study of the behaviour of imploding conical shock waves was undertaken by Skews et al [2] in which good correlation was exhibited with the patterns of reflection for waves of various strengths. This study extends that study to various incident wave strengths.

A supersonic ejector uses a primary flow expanded from high pressure to supersonic speeds to entrain a secondary flow and pump it to higher pressures. Figure1 shows the schematic of the flow through an ejector that has complex interactions between shock, shear layers, boundary layers within a varying area duct. The understanding of mixing between a supersonic stream and its coflow within a duct is crucial in design of a supersonic ejector. The length of the duct required for complete mixing of the streams followed by shock trains is specified by rule of thumb methods. Hence most of the literature deals with designing ejector for various applications and there is scarcity of experimental data on mixing phenomena in ejector, an essential requirement for optimal design considerations[3]. This motivated setting up a 2D supersonic ejector facility at the Laboratory for Hypersonic and Shock Wave Research, IISc, Bangalore, for experimentation on flow mechanics and mixing studies within the ejector. Static pressure measurements along the wall and shadowgraphs obtained at different operating conditions and their interpretations towards the objectives are described in this paper.

When a shock wave passes an interface between two fluids with different densities, the initial perturbation on the interface will grow with time, which is known as the Richtmyer-Meshkov (RM)instability [1, 2]. Due to the academic significance in interface stability, vortex dynamics and the formation mechanism of turbulence and many applications such as inertial confinement fusion and supernova explosions, much attention has been paid to the RM instability and turbulent mixing in recent decades. Many groups in different countries, such as America, France, Japan and Israel, have carried out researches respectively and obtained original achievements. Experimentswith air-SF

6

and air-heliumwere conducted in shock tube by Benjamin

et al

. [3].

Surface dielectric barrier discharges (DBD) are known to offer potential flow control techniques for low speed flows. A well known application of DBD in low speed flows is separation control as illustrated by Little et al. [1]. For high speed flows (supersonic and hypersonic flows) there is a need for deeper understanding of their nature of interaction and applicability. Literature on low speed flows suggests the body force field created by DBD, as the reason for flow alteration (Boeuf and Pitchford [2]). High speed flows, hypersonic flows in particular, are characterized by high flow kinetic energy, and thus one cannot ignore the possible energy interactions the flow can have with the regions of high temperature in the field (as may be the case with discharge).

Scramjet-powered access to space is expected to entail flight between Mach 10 and 15, along a dynamic pressure ascent trajectory of up to 100 kPa [1]. These flow conditions are characterised by total pressures of the order of gigapascals. Expansion tubes such as X2 and X3 at The University of Queensland (UQ), which add total enthalpy and total pressure to the test gas through an unsteady expansion, are the only current facilities which have the potential to achieve these very high total pressures.

An investigation was carried out into the transition between various types of reflection of converging cylindrical shock wave segments over various wedges, and the effect of the wave incidence angle and the Mach number of the incident wave on the reflection type. Such a reflection differs significantly from those of planar waves, as both the Mach number and the incident angle of the shock wave are time dependant. The transition conditions were examined in terms of transition criteria that have been suggested for planar waves in the literature.

Shockwaves are essentially non-linear waves that propagate at supersonic speeds. Such disturbances occur in steady transonic or supersonic flows, during explosions, earthquakes, hydraulic jumps and lightning. Rapid movement of piston in a tube filled with gas generates a shock wave. Any sudden release of energy (within few

μ

s) will invariably result in the formation of shock waves since they are one of the efficient mechanisms of energy dissipation observed in nature. The dissipation of mechanical, nuclear, chemical, and electrical energy in a limited space will usually result in the formation of a shock wave. Because of the dissipative nature of shock waves they invariably need a medium both for generation as well as for propagation.

A gene is a unit of heredity in a living organism. It normally resides on a stretch of DNA that codes for a type of protein or for an RNA chain that has a function in the organism. All living things depend on genes, as they specify all proteins and functional RNA chains. Genes hold the information to build and maintain an organism’s cells and pass genetic traits to offspring. The gene has to be transferred to bacteria or eukaryotic cells for basic and applied molecular biology studies. Bacteria can uptake exogenous genetic material by three ways: conjugation, transduction and transformation. Genetic material is naturally transferred to bacteria in case of conjugation and transferred through bacteriophage in transduction. Transformation is the acquisition of exogenous genetic material through cell wall. The ability of bacteria of being transformed is called competency and those bacteria which have competency are competent cells. Divalent Calcium ions can make the bacteria competent and a heat shock can cause the bacteria to uptake DNA. But the heat shock method cannot be used for all the bacteria. In electroporation, a brief electric shock with an electric field of 10-20kV/cmmakes pores in the cell wall, facilitates the DNA to enter into the bacteria. Microprecipitates, microinjection, liposomes, and biological vectors are also used to transfer polar molecules like DNA into host cells.

The starting-characteristic of inlet is one of the key factors that govern the performance of hypersonic airbreathing propulsion system. For efficient operation, the inlets must operate in a started mode. Inlet starting has been extensively studied,[1][2] however, it is difficult to accurately predict whether the inlet is started or unstarted. Therefore, it will be useful to find an easy way that is capable of testing various behaviors of the inlet starting process. Pulse facilities could play an important role in these ground tests. But it has been shown that the inlet could be started with larger internal contraction ratio (ICR) in pulse facilities,[1] such as a shock tunnel, because of the unsteady effects in flow establishment of the facility[3] which have strong capability of helping inlet to start. So, there is a large discrepancy compared with conventional facilities. However, every coin has two sides. Whether the strong help-to-start capability of pulse facility can be switched ’on’ and ’off’? The present paper reports our recent progress related to the above ideas.

Shock waves have traditionally been studied and characterized by their movement through media. And alternatively, the parameters of a medium have been successfully diagnosed by its influence on a shock-wavesmotion. The propagation of shock waves through media with various types of inhomogeneity has always been of particular interest for a wide range of application in different fields of physics, like astrophysics, laser physics, aerodynamics and many others. There are lots of experimental, theoretical and numerical studies concerning interactions of shock waves with turbulence [1], vortices [2], near-wall fine particle or dust layers [3], preshock inhomogeneities induced by non-uniform energy deposition [4], isolated bubbles [5], thermal layers [6, 7], plasma formations [8, 9], etc. These interactions are apparently mutual - not only are the dynamics and the structure of waves altered but also the properties of the medium are changed in a post shock flow. The present work studies the structure of the gas-dynamic flow resulting from the propagation of a plain shock wave along a near-wall expanding layer of the gas excited by a high-current sliding surface discharge (“plasma sheet”).

Many experiments and numerical simulations for a bow shock that forms over a blunt body have been conducted. In general, the bow shock formed in a uniform flow is stable, and a steady bow shock can be easily obtained. However, instability of the bow shock was observed in front of nearly flat bodies in a difluorodichloromethane atmosphere, using a ballistic range 30 years ago [1]. Baryshnikov et al. classified the features of this bow-shock instability into three types: small deformation (Fig. 1(a)), large deformation (Figs. 1(b) and (c)), and complete disruption of shock wave (Fig. 1(d)). From experiments under various conditions, it was concluded that bow-shock instability occurs depending on not only the Mach number and atmospheric pressure, but also the roundness of the edge and the curvature of the body surface. They suggested two candidates for the main mechanism of this phenomenon. One is dynamical nonequilibrium behind the shock wave due to a low specific heat ratio

γ

of the difluorodichloromethane; the other is chemical nonequilibrium with a quick increase in temperature at the shock front. Since direct experimental analysis of these mechanisms is difficult, numerical analysis using a sophisticated computational fluid dynamics (CFD) technique is expected to identify the mechanism that has not yet been revealed.

Experiments with blast waves at a reduced scale can be of much use to acquire a qualitative understanding of the highly complex, large scale flow phenomenon, provided appropriate scaling laws can be developed. This being the motivation, microblast waves have been generated in the laboratory using several means in the recent past. Jiang [1] had used laser to deposit energy within a small region (point source of energy), Kleine [2] and Settles [3] reported generation of micro-blast waves by ignition/ detonation of gram sized explosive charges. Interestingly, the work of Kleine [2], using milligram sized silver azide charges, had shown the validity of the popular cubic root scaling laws, hitherto used with the large scale blast waves, for the milligram charges too. This paves the way for the use of micro blast waves as an experimental tool for a qualitative study of the large scale blast waves, within the laboratory, which is impossible with the large scale blast waves.

Currently, a new generation of scientific aerospace vehicles, using advanced hypersonic airbreathing propulsion based on supersonic combustion technology, is in development at several research centers [1].

The aerospace technological products have grown that one cannot conceive of putting payloads (satellites) into Earth orbit or beyond using technologies in operation (rockets carry out solid or liquid fuel). The knowledge required to keep the current launching vehicles is already so high that if the countries do not have a technological support for their own industry, they will depend on of the supplier countries and not have independent capacity sustained. Aerospace vehicle limitations for launching payloads into orbit or beyond require a continuous reduction in size, weight and power consumption of launch vehicles. Some solutions to these challenges require paradigm shifts, new production methods, and new technologies of strategic nature. The requirements of platformslaunched satellites, high performance and reliability, as well as the strict limitations of fuel (reduction of size, weight and power consumption) for launching payloads into orbit or beyond provide the development of hypersonic aircraft using hypersonic airbreathing propulsion based on supersonic combustion.

Experimental investigations conducted in ground-based test facilities are essential for the successful development of hypersonic airbreathing propulsion based on supersonic combustion [1].

The propagation of converging polygonal shocks was studied theoretically and numerically by Schwendeman and Whitham (1987)[1]. Using the approximate theory of geometrical shock dynamics (GSD), they found solutions of the behaviour of cylindrical polygonal shock waves. They showed that an initial polygonal shape repeats at different intervals during the converging process. We have conducted experiments creating similarly shaped shock waves and compared with their work.

At high flight speeds, radiation becomes an important component of aerodynamic heat transfer, and its coupling with the flow field can significantly change the macroscopic features of the flow. As radiating flight conditions are typically encountered in re-entry trajectories, the associated flight regimes range from rarefied to continuum, and may have many levels of thermal, chemical and electronic nonequilibrium. Accurate estimates of the nonequilibrium radiation involved in high speed operations such as reentry are essential in order to more efficiently design thermal protection systems.

The cylinder of explosive with hollow cavity in one side and detonator at opposite side is called as hollow charge [1]. Chemical energy generated by initiation of explosive is focused on the center of hollow cavity. And this concentrated force generates a jet with high penetration power. This phenomenon is well known as Munroe effect, and the generated jet is called as Munroe jet. When the hollow cavity is lined thin layer of metals and various materials, the liner form jet with higher penetration power than hollow charge. This is called as shaped charge which utilizes Neumann effect. Because many studies about hollow charge and shaped charge [2] are focused on property of penetration which is one of the most important factors of these phenomena, explosive and various metals have been used in experiments. In this study, simulation experiments of Munroe jet were conducted by replacing detonator and explosive in hollow charge by impactor with a velocity up to 800 m/s against shaped gelatin [3].

This paper describes the application of a new DSMC code, called

dsmcFoam

, which has been written within the framework of the open-source computational fluid dynamics (CFD) toolbox OpenFOAM[1]. The main features of

dsmcFoam

are its C++ modularity, its unlimited parallel processing capability and its ability to easily handle arbitrary, complex 3D geometries. Results of initial benchmark trials [2] have shown excellent agreement with both analytical solutions and other conventional DSMC codes.