Oblique shock

Despite over fifty years of research on shock wave boundary layer effects and interactions, many related technical issues continue to be controversial and debated. The present survey provides an overview of the present state of knowledge on such effects and interactions, including discussions of: (i) general features of shock wave interactions, (ii) test section configurations for investigation of shock wave boundary layer interactions, (iii) origins and sources of unsteadiness associated with the interaction region, (iv) interactions which included thermal transport and convective heat transfer, and (v) shock wave interaction control investigations. Of particular interest are origins and sources of low-frequency, large-scale shock wave unsteadiness, flow physics of shock wave boundary layer interactions, and overall structure of different types of interactions. Information is also provided in regard to shock wave investigations, where heat transfer and thermal transport were import...

An experimental study of fin-generated, shock-wave, turbulent, boundary-layer interactions confirmed previous observations that, sufficiently far from the fin apex, such interactions become conical. For Mach numbers from 2.5 to 4 and fin angles from 4 to 22 deg, the inception length to conical symmetry was found to increase weakly with Mach number. Also, the inception length was found to depend primarily on the inviscid shock angle, this angle ranging from 21 to 40 deg. The inception length decreased with shock strength and was small (approximately constant at three boundary-layer thicknesses in length) at the largest shock strengths encountered. (") = normalized by 6 (~) = nondimensionalization, Eq. (3)

In this paper, the regular reflection (RR) to Mach reflection (MR) transition of asymmetric shock waves is theoretically studied by employing the classical two- and three-shock theories. Computations are conducted to evaluate the effects of expansion fans, which are inherent flow structures in asymmetric reflection of shock waves, on the RR → MR transition. Comparison shows good agreement among the theoretical, numerical and experimental results. Some discrepancies between experiment and theory reported in previous studies are also explained based on the present theoretical analysis. The advanced RR → MR transition triggered by a transverse wave is also discussed for the interaction of a hypersonic flow and a double-wedge-like geometry.

Shock wave profiles of sapphire (single-crystal Al2O3) with seven crystallographic orientations (c, d, r, n, s, g, and m-cut) were measured with time-resolved VISAR (velocity interferometer for a surface of any reflector) interferometry at shock stresses in the range 16–86 GPa. Shock propagation was in the direction normal to the surface of each cut. The angle between the c-axis of the hexagonal representation of the sapphire crystal structure and the direction of shock propagation varied from 0 for c-cut up to 90° for m-cut in the basal plane. Based on published shock-induced transparencies for three directions of shock propagation, shock-induced optical transparency correlates with the smoothness of the mechanical shock wave profile. The ultimate goal was to find the direction of shock propagation for which shock-compressed sapphire is most transparent as a window material. In the experiments particle velocity histories were recorded at the interface between a sapphire crystal and...

In a non-spherical stellar explosion, non-radial motions become important near the stellar surface. For realistic deviations from spherical symmetry, non-radial flow dramatically alters the dynamics and emission of shock emergence on a significant fraction of the surface. The breakout flash is stifled, ejecta speeds are limited, and matter is cast sideways. Non-radial ejection allows for collisions outside the star, which may engender a new type of transient. Strongly oblique breakouts are most easily produced in compact stellar progenitors, such as white dwarfs and stripped-envelope core-collapse supernovae. We study the shock structure and post-shock acceleration using conservation laws, a similarity analysis, and an approximate theory for oblique shocks. The shock is likely to extend vertically from the stellar surface, then kink before joining a deep asymptotic solution. Outflow from the region crossed by an oblique shock is probably unsteady and may affect the surface ahead of the main shock. We comment on the implications for several notable explosions in which the non-spherical dynamics described in this paper are likely to play an important role. We also briefly consider relativistic and superluminal pattern speeds.

An earlier treatment of the diffraction of a shock wave advancing into a region of uniform flow, based on Chisnell's (1965) extension of Whitham's (1957) rule for shock diffraction, is corrected for an algebraic error and then compared with an analogous treatment based on the more recent extension derived by Whitham (1968). The basis for comparison is the pressure just behind a shock wave that is diffracted by a thin wedge travelling at supersonic speed. The approximation provided by Whitham's extension is both simpler than, and typically superior to, that provided by…

Wall temperature variations significantly influence the stability and transition dynamics of hypersonic boundary layers. This study numerically investigates the stability response to thermal gradients, with wall temperatures ranging from 262 K to 362 K, for double-cone and cone-cylinder-flare (CCF) geometries. The CCF geometry, modeled using data from the Research Task Group NATO-AVT-346, consists of a 5°half-angle cone followed by a 147.3 mm cylindrical section and a 12°flare, with the sharp-nosed variant (0.1 mm nose radius) analyzed to emphasize instability mechanisms such as the first and second Mack modes. For the double cone, modeled after the ROTEX-T flight-test vehicle (457.31 mm in length, 1 mm nose radius, and 7°half-angle) at Mach 5.7, flow conditions were representative of the AFOSR-Notre Dame Large Mach 6 Quiet Wind Tunnel. Heating the entire wall stabilized second-mode instabilities while amplifying the first mode, particularly within the separation bubble. In contrast, for the CCF geometry, wall heating upstream of the separation bubble consistently stabilized both first-and second-mode instabilities throughout the flowfield. Numerical stability analyses were conducted using the JoKHeR Parabolized Stability Equations (PSE) code. These results underscore the geometry-dependent nature of thermal stabilization strategies and suggest that targeted upstream heating can effectively suppress boundary layer instabilities and delay transition in hypersonic flows.

Three-dimensional shock wave reflection comprises flow physics that is significantly different from the well-documented two-dimensional cases in a number of aspects. The most important differentiating factor is the sweep of the shock system. In particular, this work examines the nature of flow fields in which there is a transition of shock reflection configuration in three-dimensional space. The flow fields investigated have been made to exist in the absence of edge effects influencing the shock interaction and transition, as found previously to exist in conventional double-wedge studies. In general, the shock configurations are those with central regular and peripheral Mach reflection portions. It is shown that the sweep angle of the portions on either side of the transition point is subject to a cusp, as per an analytical model that is developed. This is confirmed with the use of numerical models with additional evidence provided by experimental results using oblique shadow photog...

This paper presents a two-dimensional investigation into the effectiveness of trapping shock and blast waves in a duct in order to enhance attenuation, by placing an array of opposing wedges in the channel. The concept of the wedge arrangement in the trap is to allow easy shock wave entry, with weak reflected shocks, into the trap, but stronger internal reflected shocks if a wave is re-emering. The internal reflections, including those of vortices shed from earlier shock passage, result in strong shock attenuation. Different wedge placements, wedge angles, and area blockages are investigated numerically, as well as experimentally for a particular case, using pressure measurement and schlieren photography.

Studies of shock-vortex interactions in the past have predominantly been numerical, with a number of idealizations such as assuming an isolated vortex and a plane shock wave. In the present case the vortex is generated from flow separation at a corner. A shear layer results which wraps up into a spiral vortex. The flow is impulsively initiated by the diffraction of a shock wave over the edge. The strength of the shock determines the nature of the flow at the corner and that induced behind the diffracted wave. A wide variety of cases are considered using different experimental arrangements such as having two independent shock waves arriving at the corner at different times, to reflecting the diffracting wave off different surfaces back into the vortex, and to examining the flow around bends where the reflection off the far wall reflects back onto the vortex. The majority of studies have shown that the vortex normally retains its integrity after shock transit. Some studies with curved...

This paper presents the results of a numerical study of shock/shock interactions that include both the Edney type IV and type III interactions, with emphasis on the type IV interactions. Computations are made using the direct simulation Monte Carlo (DSMC) method of Bird for Mach 10 air flow, as produced in the ONERA R5Ch low-density wind tunnel. The simulations include the flow about a shock generator which creates a relatively weak oblique shock that impinges on a much stronger cylinder bow shock. The sensitivity and characteristics of the interactions are examined by varying the horizontal distance separating the shock generator leading edge and cylinder. Results of the simulation for one separation distance are compared with wind tunnel measurements. Comparisons are made for surface heating and pressure and for flow-field values of density and rotational temperatures, as obtained with the Dual-line Coherent Anti-Stokes Scattering (DL-CARS) technique. The comparisons between exper...

Reducing wave drag associated with supersonic flight vehicles and generating vehicle steering forces using a novel flow control technique are investigated both experimentally and computationally. Experimental work involved exploring oblique shock wave attenuation for a pair of oblique shock waves generated due to a supersonic flow past a double wedge system. Computational efforts were focused on flow control over two flows of interest; 1) drag reduction on a 2D wedge surface by counter flow jets issuing from a slit located on the wedge near the leading edge, and 2) steering force generation on a missile fin by jets issuing from on surface slits located at optimum places. Three different slit configurations, at freestream Mach numbers, 0.8, 1.5, at different angles-of-attack (AoA) were studied. The aim of the work was to achieve fin lift obtained at an AoA without moving the fin but using the flow control jet. Results indicate that counterflow jet can reduce drag over a 2-D wedge up to 15%. Results also indicate that while the technique generates significant control forces for Mach=0.8 case, it fails for the Mach=1.5 flow due to flow separation over the fin surface induced by the counter flow jets.

Compressible DNS of transitional and turbulent flow in a low pressure turbine cascade 1 RAJESH RANJAN, SURESH DESHPANDE, ROD-DAM NARASIMHA, JNCASR -Direct numerical simulation (DNS) of flow in a low pressure turbine cascade at high incidence is performed using a new in-house code ANUROOP. This code solves compressible Navier-Stokes equations in conservative form using finite volume technique and uses kinetic-energy consistent scheme for the flux calculations. ANUROOP is capable of handling flow past complex geometries using hybrid grid approach (separate grid topologies for the boundary layer and rest of the blade passage). This approach offers much more control in mesh spacing and distribution compared to elliptic grid technique, which is used in many previous studies. Also, in contrast to previous studies, focus of the current work is mainly on the boundary layer flow. The flow remains laminar on the pressure side of the blade, but separates in the aft region of the suction side leading to transition. Separation bubbles formed at this region are transient in nature and we notice multiple bubbles merging and breaking in time. In the mean flow however, only one bubble is seen. Velocity profiles very near to the leading edge of the suction side suggest strong curvature effect. Higher-order boundary layer theory that includes effect of curvature is found to be necessary to characterize the flow in this region. Also, the grid convergence study reveals interesting aspects of numerics vital for accurate simulation of this kind of complex flows.

In order to be able to judge the effectiveness of transition induction in WP-2, reference flow cases were planned in WP-1. There are two obvious reference cases—a fully laminar interaction and a fully turbulent interaction. Here it should be explained that the terms “laminar” and “turbulent” interaction refer to the boundary layer state at the beginning of interaction only. There are two basic configurations of shock wave boundary layer interaction and these are a part of the TFAST project. One is the normal shock wave, which typically appears at the transonic wing and on the turbine cascade. The characteristic incipient separation Mach number range is about M = 1.2 in the case of a laminar boundary layer and about M = 1.32 in the case of turbulent boundary layer. The second typical flow case is the oblique shock wave reflection. The most characteristic case in European research is connected to the 6th FP IP HISAC project concerning a supersonic business jet. The design speed of thi...

Mechanisms of the particle acceleration at a collisionless shock have been intensively studied analytically, numerically, and observationally. Most of the previous studies assume that energetic particles interact with a single shock. However, shock waves are ubiquitous in space, and two shocks frequently come close to or even collide with each other. For instance, it is observed that a CME (coronal mass ejection) driven shock collides with the earth's bow shock [H. Hietala et al., 2011], or interplanetary shocks pass through the heliospheric termination shock [J. Y. Lu et al., 1999]. The detailed structures of such colliding shocks and the accompanied particle heating/acceleration processes have not been understood. Cargill et al. [1986] performed one dimensional hybrid simulations to discuss the dynamic structure of colliding shocks and the accompanied ion acceleration. They showed that some ions are efficiently accelerated at the time of the collision of two supercritical shocks. However, since electron dynamics are neglected in a hybrid simulation, the microstructures of the colliding shocks, which may affect the early stage processes of particle acceleration, cannot be resolved. Here, we perform full Particle-in-Cell (PIC) simulations to examine colliding two shocks. In particular, the following three points interacting with two colliding oblique shocks is discussed in detail. 1. Energetic electrons are observed upstream of the two shocks before their collision. These energetic electrons are efficiently accelerated through multiple reflections at the two shocks (Fermi acceleration). Moreover, a part of the accelerated electrons are farther energized by interacting with increasing magnetic field during the collision and/or one of the shocks after the collision. 2. Before two shocks collide, there is a large amplitude wave excited by electrons flowing out to the upstream. We discuss the excitation mechanism and the influence on particle propagation or shock structures. 3. After two shocks collide, we find that a plasma density and pressure in the downstream is lower than the value calculated by MHD. The reason is that energetic electrons run away to the upstream. In addition to this, we discuss kinetic influences by comparing PIC simulation's results with the value after two shocks collide (magnetic fields, the shock velocity and etc.) calculated by MHD.

The interaction of an oblique shock wave and a laminar boundary layer developing over a flat plate is investigated by means of numerical simulation and global linear-stability analysis. Under the selected flow conditions (free-stream Mach numbers, Reynolds numbers and shock-wave angles), the incoming boundary layer undergoes separation due to the adverse pressure gradient. For a wide range of flow parameters, the oblique shock wave/boundary-layer interaction (OSWBLI) is seen to be globally stable. We show that the onset of two-dimensional large-scale structures is generated by selective noise amplification that is described for each frequency, in a linear framework, by wave-packet trains composed of several global modes. A detailed analysis of both the eigenspectrum and eigenfunctions gives some insight into the relationship between spatial scales (shape and localization) and frequencies. In particular, OSWBLI exhibits a universal behaviour. The lowest frequencies correspond to stru...

Electric field measurements at oblique, subcritical bow shock crossings are presented from the ISEE 1 University of California, Berkeley, double‐probe electric field experiment. The measurements averaged over the 3‐s spin period of the spacecraft provide the first observations of the large‐scale (100 km) laminar oscillations in the longitudinal component of the electric field associated with the whistler precursor which is characteristic of these dispersive shocks. The amplitude of the oscillations increases from ∼0.5 mV/m to a maximum of 6 mV/m across the magnetic ramp of the shock (directed along the shock normal). The calculated electric potential drops across the shocks varied from 340 to 550 volts, which is 40–60% of the observed loss of kinetic energy associated with the bulk flow of the ions. These measurements suggest that at these shocks the additional deceleration of incident ions is due to the Lorentz force. The contributions to the normal component of the large‐scale ele...

The interaction of a plane shock wave in air with concave profiles has been used in the past mainly to understand the nature of shock wave focusing. The current study examines the complex two-dimensional flow field resulting from the interaction of a plane shock wave entering a symmetrical cavity with curved walls. Of particular interest are the development of reflection patterns of the incident shock wave at the profile wall and the process of gas dynamic focus. These principal flow features are examined across a wide range of different reflector shapes. This includes a review of previously studied profiles such as cylindrical and parabolic, and also of a number of additional profiles, including compound profiles, where an inlet profile merging with that of the main cavity is shown to have major effects on the focusing mechanism and pressures. The various reflector shapes were specified by varying the shape of the profile and the depth-to-aperture ratio. The strength of the incident plane shock wave was limited between Mach numbers of 1.04 and 1.45. The principal flow features were established and examined experimentally using a variety of qualitative and quantitative flow visualization techniques, supplemented with numerical results. Time-resolved high-speed imaging was used to capture the interaction providing the unique ability to track the various transient flow features over the course of the interaction. The three primary factors that influence the maximum pressure amplification at focus, and the focus mechanism, are the incident shock strength, the depth-to-aperture ratio of the profile and an inlet profile leading into the main cavity, if present. An inlet profile results in higher-pressure amplifications for corresponding shock strengths and depth-to-aperture ratios. Increases in the depth-to-aperture ratio increase the maximum pressure amplification observed at focus. This occurs due to a combination of factors including: the strengthening of the individual shock waves involved in focus; the duration of focus; and the strengthening of a compressive flow field that develops adjacent to the shock system during focus. The compressive flow field adjacent to the shock system at focus is shown to be of great importance to the focus process.

The factor which is of prime importance in influencing the shock reflection geometry, and resulting pressures, following impingement of a shock wave on a porous surface is the velocity of the flow into the surface. A set of experiments has been conducted, using holographic inferometry in a shock tube, on the impingement of a shock wave on a surface covered with slits, over the full range of shock incidence angles from 0 to 90°. Inverse shock pressure ratios of 0.4, 0.5 and 0.7 were used, and detailed characterization of the flow fields determined. A number of methods are used to infer the inflow into the surface, and measurements are also conducted on the downstream side of the slit plate in order to establish the pressure ratio across the plate. The tests include choking of the flow through the slits. Shock reflection angles are found to be depressed compared to reflection from an impervious wall for cases of regular reflection, but are similar in the case of Mach reflection with t...

Previous detailed studies of the interaction of a shock wave with a perforated sheet considered the impact of a shock wave on a plate with regularly spaced slits giving area blockages of 60 and 67%, at various angles of incidence, and resulting in both regular and Mach reflection. The current work extends this study to a much wider variety of plate geometries. Blockage ratios of 20, 25, 33, 50, and 67 and inclinations of 45, 60, 75, and 90°to the shock wave were tested. Four different thicknesses of plate were tested at the same frontal blockage in order to assess the effects of gap guidance. Tests were conducted at two shock Mach numbers of 1.36 and 1.51 (inverse pressure ratios of 0.4 and 0.5). It is found that secondary reflected and transmitted waves appear due to the complex interactions within the grid gaps, and that the vortex pattern which is generated under the plate is also complex due to these interactions. The angle of the reflected shock, measured relative to the plate, decreases with plate blockage and the angle of inflow to the plate reduces with increasing blockage. By analysing the flow on the underside of the plate the pseudo-steady flow assumption is found to be a reasonable approximation. Both the pressure difference and the stagnation pressure loss across the plate are evaluated. It is found that over the range tested the plate thickness has a minimal effect.

A comprehensive numerical study was carried out to investigate the unsteady cell-like structures of oblique detonation waves (ODWs) for a fixed Mach 7 inlet flow over a wedge of 30°turning angle. The effects of grid resolution and activation energy were examined systematically at a dimensionless heat addition of 10. The ODW front remains stable for a low activation energy regardless of grid resolution, but becomes unstable for a high activation energy featuring a cell-like wave front structure. Similar to the situation with an ordinary normal detonation wave (NDW), a continuous increase in the activation energy eventually causes the wave-front oscillation to transit from a regular to an irregular pattern. The wave structure of an unstable ODW, however, differs considerably from that of a NDW. Under the present flow condition, triple points and transverse waves propagate downstream, and the numerical smoke-foil record exhibits traces of triple points that rarely intersect with each other. Several instability-driving mechanisms were conjectured from the highly refined results. Since the reaction front behind a shock wave can be easily destabilized by disturbance inherent in the flowfield, the ODW front becomes unstable and displays cell-like structures due to the local pressure oscillations and/or the reflected shock waves originating from the triple points. The combined effects of various instability sources give rise to a highly unstable and complex flow structure behind an unstable ODW front.

The supersonic combustion scramjet in the inlet applies the shock waves compression mechanism tosubstitute the actual compressor from a gas turbine engine. The scramjet works with combustion of fuel throughthe air stream in supersonic condition at least with Mach 5. Novel design of a scramjet intake system was madewith variations in the angle of the fins and entrance width. The best combination of diameter and inclinationangle was 1.75 m and 15 degrees, respectively. The findings were able to increase the oblique shock waveinteractions and supplicate effective combustion and reduce pressure losses for the effective application ofscramjet system, which can be significant for aerospace industry.

The capability of the NavierStokes equations with a perfect gas model for simulation of hypersonic shock wave laminar boundary layer interactions is assessed. The con¦guration is a hollow cylinder §are. The experimental data were obtained by Calspan-University of Bu¨alo (CUBRC) for total enthalpies ranging from 5.07 to 21.85 MJ/kg. Comparison of the computed and experimental surface pressure and heat transfer is performed and the computed §ow¦eld structure is analyzed.

In this paper, by numerical simulation of gas flow, effects of divergence angles on the amplification factor, side force and thrust vector magnitude of a moveable nozzle with supersonic split line have been investigated. The study is conducted for constant exit to throat area ratio and the location of the split line is also constant. The finite volume method with second order upwind scheme for spatial discretization and AUSM method for flux splitting have been used in numerical solution of the equations. It was observed that for two different deviation angles, the results of RNG k-ε turbulence model are in good agreement with experimental data. Results also show that by increasing the divergence angle of the nozzle, the amplification factor, side force and thrust vector magnitude will decrease.

Unsteadiness of axisymmetric shock-dominated hypersonic laminar separated flow over a double cone is studied for the first time using a combination of time accurate Direct Simulation Monte Carlo (DSMC) calculations, linear global instability analysis, and momentum potential theory (MPT). Close to steady state linear analysis reveals the spatial structure of the underlying temporally stable global modes. At all Reynolds numbers examined, the amplitude functions demonstrate the strong coupling between the separated flow region at the cone junction with the entire shock system, including pressure and temperature waves generated behind the shock and spatially amplified Kelvin-Helmholtz waves. In addition, as the Reynolds number is increased, temporally damped harmonic shock oscillations and multiple-reflected λ-shock patterns emerge in the eigenfunctions. Application of the MPT (valid for both linear and nonlinear signals) to the highest Reynolds number DSMC results shows that large aco...

The effects of overtaking disturbances behind the flow on the propagation of diverging cylindrical shock Waves through an ideal gas in presence of a magnetic field having =constant=  and an Initial density distribution  where  is a constant,  is the density at the plane / exes of symmetry: The analytical formula for flow variables representing both the position form viz; weak and strong cases at shock waves have been obtained. Their numerical estimates at permissible shock front locations have been obtained. There numerical estimates at permissible shock front location's have been Calculated and compared with earlier result describing in Free Propagation through figures. After inclusion of E.O.D. noted that there is no change at flow variable with parameters and . However, the trends of variation with propagation distance r, for shock strength, shock velocity and particle velocity are not change in case of weak shock with work Magnetic field(wswmf).

and Amagats laws are two well-known thermodynamic models describing gas mixtures. Our current research is focused on determining the suitability of these models in predicting effects of shock propagation through gas mixtures. Experiments are conducted at the Shock Tube Facility at the University of New Mexico (UNM). The gas mixture used in these experiments consists of approximately 50% sulfur hexafluoride (SF6) and 50% helium (He) by mole. Fast response pressure transducers are used to obtain pressure readings both before and after the shock wave; these data are then used to determine the velocity of the shock wave. Temperature readings are obtained using an ultra-fast mercury cadmium telluride (MCT) infrared (IR) detector, with a response time on the order of nanoseconds. Coupled with a stabilized broadband infrared light source (operating at 1500 K), the detector provides pre-and post-shock line-of-sight readings of average temperature within the shock tube, which are used to determine the speed of sound in the gas mixture. Paired with the velocity of the shock wave, this information allows us to determine the Mach number. These experimental results are compared with theoretical predictions of Daltons and Amagats laws to determine which one is more suitable.

aerothermal stresses. An understanding of swept (3D) interactions is lacking compared to their unswept (2D) counterparts. To this end, the effect of disturbances induced by a microramp at the root region of a swept SBLI is investigated. Experiments are conducted at a nominal Mach number of 2.3 with a fully turbulent boundary layer (Re θ = 5.5 × 10 3). Previous investigations show unsteady shock motion with constant frequencies across the spanwise domain. To determine if this frequency content is associated with the root structure, the length scale of root separation is altered by installing a microramp of height 0.3δ 0 , placed 15δ 0 upstream of the root inviscid shock impingement location. Preliminary oil flow visualization shows no significant differences on the mean flow topology away from the root, enabling direct comparisons between the baseline and perturbed unsteady behavior. Mean and unsteady wall pressures are captured to assess changes in pressure distribution and associated spectral content, gaining valuable insight into the underlying physics.

Numerical investigations were performed with a supersonic inlet system installed with a three-dimensional bump which was substituted for a diverter or conventional ramp-type compression systems at Mach 2. The modified inlets were designed to have two oblique shocks and a terminal normal shock followed by a subsonic diffuser, with a circular cross-section throughout. A numerical analysis was conducted to understand the three-dimensional flow field including shock/boundary layer interactions that occur around a three-dimensional bump and to evaluate the performance of the supersonic inlets. The current numerical simulations showed a bump-type inlet based on a conventional ramp-type inlet can provide an improvement in total pressure recovery downstream of the shock/boundary-layer interaction over a ramp-type inlet.

In the starting phases of continuously blowing under-expanded jets, this numerical study investigates the effect of co-flow (Ua U j), (a) on the circulation & evolution of primary vortex ring (PVR) and (b) on the occurrence of Mach reflection, slip-stream generation and subsequent formation of counter rotating vortex rings (CRVRs). With increase in co-flow (Ua U j), the PVR circulation gradually decreases. The size of supersonic PVR gradually decreases with increase in co-flow (Ua U j) and at high magnitudes of co-flow (Ua U j 0.3), the supersonic PVR attains a circular shape. The strengths of embedded shock (ES) and vortex induced shock (VIS) are found to decrease with increase in co-flow (Ua U j) and at high magnitudes of co-flow (Ua U j 0.3) these shocks may even cease to form inside the supersonic PVR. Increase in co-flow (Ua U j) causes the expansion fan to become more and more narrow. This reduces the acceleration of the supersonic flow inside the inviscid core, thereby weakening the incident oblique shock (IOS), which in turn increases the pressure prevailing downstream of this shock inside the inviscid core. The increase in co-flow (Ua U j) also leads to a simultaneous decrease in the pressures prevailing infront of the downstream marching PVR and Mach disk (MD) of the inviscid core due to the reduction in the strength of precursor shock. As the magnitudes of pressures prevailing in the upstream and downstream of Mach disk (MD) approach each other, hence, MD also weakens. This shows that with increase in co-flow (Ua U j), there is weakening of the different shocks (i.e. ES, IOS, MD) involved in Mach reflection. This causes a reduction in the strength of the resulting slip-stream, thereby affecting the formation of CRVR patterns.

In studies on instabilities of flowfield in rotating detonation, one of the most common concerns is the instability at the slip line originating from the conjunction of the detonation wave and oblique shock. Using Euler equations associated with 7-species-and-8-reaction finite-rate chemical reaction model of hydrogen/air mixtures, further studies are performed to simulate the 2-D rotating detonation, and the flow mechanism of instability at the slip line is investigated in depth. The results show that the distinct wake profile exists at the slip line, which is different from the typical mixing layer. Analysis indicates that the generation of wake is caused by the transition shock between the detonation wave and oblique shock. Because of the wake profile, the vorticity distribution therein appears in a double-layer layout, and different evolutions exist in different vorticity layers. Based on the veloc ity profile across the slip line, the analysis by the linear stability theory is made, and two unstable modes which have different shape profiles and phase velocities are found. Discrete Fourier transformation is utilized to analyze the numerical results, and similar shape profiles are obtained. A general coincidence in velocity of vortex movement is also attained between the theoretical predictions and simulations. Investigations show that the wake instability is responsible for the unstable mechanism, and corresponding unstable structures differs from the canonical ones in typical mixing layers.

In studies on instabilities of flowfield in rotating detonation, one of the most common concerns is the instability at the slip line originating from the conjunction of the detonation wave and oblique shock. Using Euler equations associated with 7-species-and-8-reaction finite-rate chemical reaction model of hydrogen/air mixtures, further studies are performed to simulate the 2-D rotating detonation, and the flow mechanism of instability at the slip line is investigated in depth. The results show that the distinct wake profile exists at the slip line, which is different from the typical mixing layer. Analysis indicates that the generation of wake is caused by the transition shock between the detonation wave and oblique shock. Because of the wake profile, the vorticity distribution therein appears in a double-layer layout, and different evolutions exist in different vorticity layers. Based on the veloc ity profile across the slip line, the analysis by the linear stability theory is made, and two unstable modes which have different shape profiles and phase velocities are found. Discrete Fourier transformation is utilized to analyze the numerical results, and similar shape profiles are obtained. A general coincidence in velocity of vortex movement is also attained between the theoretical predictions and simulations. Investigations show that the wake instability is responsible for the unstable mechanism, and corresponding unstable structures differs from the canonical ones in typical mixing layers.

The supersonic combustion scramjet in the inlet applies the shock waves compression mechanism to substitute the actual compressor from a gas turbine engine. The scramjet works with combustion of fuel through the air stream in supersonic condition at least with Mach 5. Novel design of a scramjet intake system was made with variations in the angle of the fins and entrance width. The best combination of diameter and inclination angle was 1.75 m and 15 degrees, respectively. The findings were able to increase the oblique shock wave interactions and supplicate effective combustion and reduce pressure losses for the effective application of scramjet system.

The supersonic combustion scramjet in the inlet applies the shock waves compression mechanism tosubstitute the actual compressor from a gas turbine engine. The scramjet works with combustion of fuel throughthe air stream in supersonic condition at least with Mach 5. Novel design of a scramjet intake system was madewith variations in the angle of the fins and entrance width. The best combination of diameter and inclinationangle was 1.75 m and 15 degrees, respectively. The findings were able to increase the oblique shock waveinteractions and supplicate effective combustion and reduce pressure losses for the effective application ofscramjet system, which can be significant for aerospace industry.

Mixing between the injected fuel and high speed free stream air is challenging at supersonic speeds. Placing cavities downstream of injection holes or slots addresses the problem of flame holding and stabilisation, however there are still open questions related to mixing enhancement, uniformity and efficiency. The present study examines experimentally the flow field interactions due to a transverse jet-cavity combination with shock impingement at supersonic speeds using PIV, Schlieren photography, and oil flow surface visualisation. The oblique shock lifts the shear layer over the cavity and combined with the instabilities generated by the transverse jet injection creates a highly complicated flowfield with numerous vortical structures. The interaction between the oblique shock and the jet leads to a relatively uniform velocity distribution within the cavity. The lifting of the shear layer is also believed to reduce the drag created by the cavity.

As a preliminary to understanding how a baryon-quark phase transition would affect the flow of nuclear matter in relativistic heavy ion collisions, we study how such a phase transition would affect the relativistic flow of nuclear matter through an oblique shock wave.

As a preliminary to understanding how a baryon-quark phase transition would affect the flow of nuclear matter in relativistic heavy ion collisions, we study how such a phase transition would affect the relativistic flow of nuclear matter through an oblique shock wave.

A theoretical model of shock-wave propagation has been studied in a heat-conducting and a self-gravitating medium. The effects of magnetic field has been taken into consideration. The shock is strong enough to neglect the ambient gas pressure. The variation of flow variables behind the shock have been investigated numerically.

In the present research the characteristics of nonequilibrium flow are studied by applying CARS (Coherent Anti-Stokes Raman Spectroscopy) method to hypervelocity shock waves in low-density air. Furthermore, in order to get more detail information on the flow behind the hypervelocity shock wave and to inspect whether the disturbance is seen at the shock wave front during the CARS measurement, radiation images behind the shock waves are obtained simultaneously. Finally, the CARS signals are detected behind hypersonic shock waves with a shock velocity of up to 7km/s. In addition, the radiation distributions are obtained at the same experiment. The vibrational and rotational temperatures are estimated by spectral fitting method. It is found that the rotational temperature exceeds the vibrational temperature when shock wave velocity is 5km/s or more, where both temperatures remains lower than translational temperature.

The purpose of the present study was to investigate the interaction between stream-wise vortices with oblique shock fronts of various intensities. A three-dimensional, Navier-Stokes solution of the problem with an orthogonal grid incorporating a RANS approach with a RNG kturbulence model was used to simulate earlier experimental studies of the problem. The accuracy of the computational study was examined by comparing the results to the previous experiments of oblique shock vortex interactions using three different shock wave intensities leading to weak, moderate and strong interactions as identified in our previous experimental investigations. Highlights Developed a numerical model for the Interaction between supersonic wing-tip vortices and oblique shock waves The Strong Interaction led to Vortex Breakdown correctly modelled by the numerical scheme and the RNG k-ε turbulence model Experimental distributions of properties of the vortex were used to reproduce the vortical flow at the inlet Results were validated against previous experimental activities and compared to a relevant previous work of Rizzetta [28]