Unsteady Shock Wave Research Articles

Interaction of an Optical Discharge with a Shock Wave

The results of an experimental study of the impact of the focused pulsed-periodic radiation from a CO2 laser on a gas-dynamic structure in a supersonic jet are presented. The radiation of the CO2 laser is propagated across the stream and focused by a lens on the axis of the supersonic jet. To register the flow structure, a shadow scheme with a slit and a flat knife located along the flow is used. The image is fixed by a speed camera with an exposure time of 1.5 μs and a frame rate of 1000 1/s. In the flow, the plasma initiated by the pulsedperiodic laser is visualized in order to identify and determine the period of plasma development, as well as the motion of the initial front of the shock wave. It is shown that at the transverse input of laser radiation into the stream the periodic structure of the thermal trace is created with the formation of an unsteady shock wave from the energy release zone. At small repetition rates of laser radiation pulses, the thermal spot interacts with the flow in the pulsed mode. It is shown that elliptic nonstationary shock waves are formed only at low subsonic flow velocities and in a stationary atmosphere. The process of nonstationary ignition by an optical discharge of a methane–air mixture during a subsonic outflow into a motionless atmosphere is shown experimentally. The results of optical visualization indicate burning in the trace behind the optical discharge region.

Numerical Simulation of Rotor in Hover and Forward Flight

The flow around rotor is numerical simulated in hover and forward flight based on multi-structured grid. In hover the flow field can be transformed into a steady-state flow field in the rotating coordinate system. The experimental data of Caradonna and Tung rotor is used to verify the numerical simulation result. The numerical results compare well with the experimental data for both non-lifting and lifting cases. Non-lifting forward flight is simulated and the prediction capabilities have been validated through the ONERA two-blade rotor. The pressure distributions of different positions under different azimuth angles are compared, which is in good agreement with the experimental data. There is unsteady shock wave when forward flight. Dual-time method is used to obtain unsteady flow field with rigid moving grid in the inertial system.

Tracking background-oriented schlieren for observing shock oscillations of transonic flying objects.

A detailed understanding of the in-flight behavior of high-speed flying objects is useful in the forensic investigation of crime scenes. In particular, for transonic flying objects with associated unsteady shock waves, long-duration and high-resolution measurements are desirable but difficult with conventional optical visualization methods. In this study, we propose a tracking background-oriented schlieren (BOS) method that uses a mirror-based high-speed optical axis controller and a striped retroreflective background. We combine prediction instruction including a compensator for the flight trajectory, high-speed visual feedback by silhouette recognition, and reference image generation from only a measurement image in the BOS visualization process. Experimentally, we successfully observe the shock oscillations of an actual high-speed flying object.

Energy budget analysis of aeroelastic limit-cycle oscillations

In this work limit-cycle oscillations (LCOs) caused by aerodynamic non-linearities were analysed using fluid-structure interaction (FSI) simulations of a two-degree-of-freedom airfoil system. A comparison with sinusoidal forced motion oscillation simulations at the fundamental frequency was performed. The results show that using the first harmonic component for the forced motion oscillations is a good approach to describe the limit-cycle oscillations considered in this work. Following the work of G. Dietz, G. Schewe and H. Mai (2004), Journal of Fluids and Structures, Vol. 19, an energy budget analysis of the limit-cycle oscillations was performed. Power analysis of the LCO reveals that the power of the lift and moment show non-linear behaviour with increasing amplitude. The linearised equivalent power components that would occur in case of flutter were computed. These show that the defect in the power of the lift in the non-linear case is caused by the increase of the phase of the lift with oscillation amplitude, which is the result of the unsteady shock wave motions on both upper and lower surface of the airfoil. The power of the moment also shows a defect, which is much smaller than in case of the lift. This defect is caused by variations in both the magnitude and the phase of the moment with oscillation amplitude.

Normal Reflection Characteristics of One-Dimensional Unsteady Flow Shock Waves on Rigid Walls from Pulse Discharge in Water

Strong shock waves can be generated by pulse discharge in water, and the characteristics due to the shock wave normal reflection from rigid walls have important significance to many fields, such as industrial production and defense construction. This paper investigates the effects of hydrostatic pressures and perturbation of wave source (i.e., charging voltage) on normal reflection of one-dimensional unsteady flow shock waves. Basic properties of the incidence and reflection waves were analyzed theoretically and experimentally to identify the reflection mechanisms and hence the influencing factors and characteristics. The results indicated that increased perturbation (i.e., charging voltage) leads to increased peak pressure and velocity of the reflected shock wave, whereas increased hydrostatic pressure obviously inhibited superposition of the reflection waves close to the rigid wall. The perturbation of wave source influence on the reflected wave was much lower than that on the incident wave, while the hydrostatic pressure obviously affected both incident and reflection waves. The reflection wave from the rigid wall in water exhibited the characteristics of a weak shock wave, and with increased hydrostatic pressure, these weak shock wave characteristics became more obvious.

Study on drug powder acceleration in a micro shock tube

Recently, micro shock tubes have been widely used in the medical engineering. The needle-free drug delivery device which mainly consists of a micro shock tube and an expanded nozzle has been produced to inject drug powders into human and animal bodies without any sharp metal needles. The drug powders were delivered by obtaining high momentum, which can be done by accelerating drug powders in the micro shock tube and supersonic nozzle. The particle-gas flows are induced by the incident shock wave developing by rupturing the diaphragm in the micro shock tube and again accelerated in the supersonic nozzle. The momentum of injected drug particles should be strictly controlled otherwise patients will suffer from skin injury or hurt. Even though micro shock tubes have been investigated in the past several decades, the detailed studies on particle-gas flows in the micro shock tube were rare to date due to the micro size and difficult experimental operation on micro shock tubes. In this paper, the experimental and numerical studies were carried out on investigating particle-gas flows in a designed micro shock tube. Particle tracking velocimetry (PTV) was performed to calculated particle average velocity at the exit of the supersonic nozzle. The nozzle flows were analyzed by obtaining instantaneous particle fields. The particle number density ratio was also investigated in the test section. The numerical simulations were performed by calculating unsteady Naver-Stokes equations on compressible flows and using fully implicit finite volume schemes. Discrete phase model (DPM) was used for simulating particle-gas flows in the micro shock tube. Particle diameter and density were varied to investigate their effects on the particle-gas flows. Unsteady particle-gas flows and shock wave propagation were obtained in details in the micro shock tube for present experimental and numerical studies.

Unsteady transonic flow control around an airfoil in a channel

Transonic internal flow around an airfoil is associated with self-excited unsteady shock wave oscillation. This unsteady phenomenon generates buffet, high speed impulsive noise, non-synchronous vibration, high cycle fatigue failure and so on. Present study investigates the effectiveness of perforated cavity to control this unsteady flow field. The cavity has been incorporated on the airfoil surface. The degree of perforation of the cavity is kept constant as 30%. However, the number of openings (perforation) at the cavity upper wall has been varied. Results showed that this passive control reduces the strength of shock wave compared to that of baseline airfoil. As a result, the intensity of shock wave/boundary layer interaction and the root mean square (RMS) of pressure oscillation around the airfoil have been reduced with the control method.

SHOCK EXCURSION IN TRANSONIC CONVEX-CORNER FLOWS

A typical feature on a transonic airfoil is associated with a quasi-normal shock on the upper surface. When a shock is strong enough, development of a separation bubble changes the flowfield significantly. High intensity of pressure fluctuations is associated with unsteady shock wave. An experimental study was conducted to investigate transonic convex-corner flows, including single- and round-convex corners. Peak pressure fluctuations and zero-crossing frequency are associated with incoming Mach number, total turning angle, and separation length. In the cases of R100 and R200, the level of zero-crossing frequency decreases with increasing turning angle as observed in the single convex-corner flows. An increase in the level of pressure fluctuations corresponds to lower zero-cross frequency. In the cases of R300, there is a roughly constant level of surface pressure fluctuations or the variation of pressure fluctuations with the turning angle is less significant. The zero-crossing frequency would be independent of turning angle in the cases of R300.

Shock-wave processes in wire explosion in water and bubbly media

The structure of unsteady shock waves and bubble detonation waves generated by electrical explosion of a wire in water and chemically inert and reactive bubbly media was studied at a capacitor energy storage of 8–81 J. The formation and expansion of a plasma bubble after wire explosion in water and bubble media was investigated using optical recording. The formation of cavitation zones and the breakdown characteristics of the liquid and bubbly media in short strong shock waves were established. It is proved that the initiation of bubble detonation by wire explosion occurs by a resonance shock-wave mechanism. A comparative frequency Fourier analysis of shock waves in an inert bubble medium and bubble detonation waves was performed.

Computation of transonic internal flow around a biconvex airfoil with cavity

At transonic flow conditions, unsteady self-excited shock waves are frequently observed in several modern internal aeronautical applications such as in turbine cascades, compressor blades, butterfly valves, fans, nozzles, diffusers and so on. The appearance of shock oscillation often causes serious problems such as aero-acoustic noise, non-synchronous vibration (NSV), intense drag rise, high cycle fatigue failure (HCF) and buffeting. In recent years, the effect of various passive means on the airfoil are investigated both experimentally and numerically to find the effectiveness as a shock control device. In the present study, the transonic internal flow around an airfoil using a cavity as a passive means of shock control was investigated numerically. Computational results are validated with available experimental data. Results showed that the airfoil with cavity significantly reduced the flow field unsteadiness such as the amplitude of pressure oscillation and root mean square of pressure oscillation.

Dynamic transition from Mach to regular reflection of shock waves in a steady flow

AbstractThe steady, two-dimensional transition criteria between regular and Mach reflection are well established. Little has been done on the dynamic effect on transition due to a rapidly rotating wedge. Results from experiments and computations done on steady and unsteady shock wave transition from Mach reflection to regular reflection, MR $\def \xmlpi #1{}\def \mathsfbi #1{\boldsymbol {\mathsf {#1}}}\let \le =\leqslant \let \leq =\leqslant \let \ge =\geqslant \let \geq =\geqslant \def \Pr {\mathit {Pr}}\def \Fr {\mathit {Fr}}\def \Rey {\mathit {Re}}\rightarrow $ RR, are described. The measured motion and the initial shock incidence was used to mimic the experiment with a two-dimensional numerical code. The maximum rotation speed achieved at transition was approximately $2500^\circ \ {\mathrm{s}}^{-1}$. Rapid wedge rotation was shown to have a significant measurable effect on transition. The code was also applied to the dependence of dynamic MR $\rightarrow $ RR transition on other variables in the parameter space. These include rotation about the leading or trailing edge, initial incidence and rotation speed at two free-stream conditions. Impulsively started rotation in these cases was used with the rotation specified by $M_E = \omega c/a_{\infty }$ where $\omega $ is constant angular velocity (negative anticlockwise), $c$ the distance from the edge considered to the pivot point and $a_{\infty }$ the free-stream sound speed. For the Mach numbers and range of rotation speeds tested, both the wedge and shock angle at transition decreased with increased rotation speed. The sensitivity of the transition angle to changing the rotation point from the trailing edge to the experimental model pivot point was investigated briefly at a free-stream Mach number of $M = 2.98$ with $M_E = -0.1$. The wedge angle at transition increased by 1.5° and the shock angle at transition decreased by 1.5°, a significant variation. The effect of the initial incidence was also investigated. By reducing the initial wedge angle from 24.5 to 23.5° for these initial conditions the shock angle at transition decreased by approximately 1.8°, also a marked sensitivity. The flow field development for impulsive rotation about the wedge trailing and leading edges at $M = 1.93$ for $M_E = -0.075$ was analysed in some detail. The flow field development is very sensitive to the rotation centre, more especially at large rotation rates. Four phases of the Mach stem development were identified in both cases. For rotation about the wedge leading edge the Mach stem height remains constant until the expansion waves arrive at the triple point. This is followed by an increase in Mach stem height, which then remains constant for a short period after which it decreases until transition to RR. For rotation about the wedge trailing edge the impulsive start generates a disturbance on the incident wave which propagates down the wave, through the triple point and down the Mach stem. The stem height is constant until the arrival of the incident wave disturbance. This causes a sudden decrease in Mach stem height. Subsequently, the Mach stem height remains constant for a short time, before it decreases until transition to RR. Similar effects in the variation of stem height with wedge angle occur at the higher Mach number of 2.98. It was demonstrated that MR can be maintained for a while at zero wedge incidence with a sufficiently large rotation rate of $M_E = -0.1$, with $M=1.93$, for both leading and trailing edge pivot points.

1406 PIV法を用いた衝撃波と渦の干渉に関する可視化計測

Interaction of unsteady shock wave and Karman's vortex is experimentally investigated in a wind tunnel. Karman's vortex is generated behind a cylindrical bar and the Reynolds number based on the mean velocity and the cylinder diameter is set to 849. A spherical shock wave is emanated from an open end of a shock tube installed far downstream of the bar and propagates in the upstream direction. The shock wave is successfully captured by a high-speed camera and instantaneous fluid velocity interacted with the shock wave is calculated by means of particle image velocimetry (PIV). The visualized velocity and vorticity distributions show that both the streamwise velocity fluctuation and vorticity of the Karman's vortex are decreased by the shock wave.

PREDICTION OF TRANSONIC BUFFET ONSET FOR FLOW OVER A SUPERCRITICAL AIRFOIL- A NUMERICAL INVESTIGATION

Transonic flow over a supercritical airfoil leads to the appearances of unsteady shock waves in theflow field. At certain flow conditions, the interaction of unsteady shock waves with boundary layer becomescomplex and generates self-excited shock oscillation, lift fluctuation and thus initiate the buffet. In the presentstudy, Reynolds averaged Navier-Stokes equations with k-? SST turbulence model has been applied to predictthe shock induced buffet onset for the flow over a supercritical airfoil NASA SC(2) 0714. The free streamtransonic Mach number is kept in the range of 0.71 to 0.75 while the angle of attack is varied in a wide range.The onset of buffet is confirmed by the fluctuating aerodynamic properties such as lift-coefficient, pressurecoefficient, static pressure and so on. The self-excited shock oscillation and the corresponding buffet frequencyare numerically analyzed.DOI: http://dx.doi.org/10.3329/jme.v43i1.15782

Method for Optimally Controlling Unsteady Shock Strength in One Dimension

This paper presents a new formulation and computational solution of an optimal control problem concerning unsteady shock wave attenuation. The adjoint system of equations for the unsteady Euler system in one dimension is derived and used in an adjoint-based solution procedure for the optimal control. A novel algorithm is used to satisfy all necessary first-order optimality conditions while locally minimizing an appropriate cost functional. Distributed control solutions with certain physical constraints are calculated for attenuating blast waves similar to those generated by ignition overpressure from the shuttle’s solid rocket booster during launch. Results are presented for attenuating shocks traveling at Mach 1.5 and 3.5 down to 85%, 80%, and 75% of the uncontrolled wave’s driving pressure. The control solutions give insight into the magnitude and location of energy dissipation necessary to decrease a given blast wave’s overpressure to a set target level over a given spatial domain while using only as much control as needed. The solution procedure is sufficiently flexible such that it can be used to solve other optimal control problems constrained by partial differential equations that admit discontinuities and have fixed initial data and free final data at a free final time.

Three-dimensional laser interferometric CT (LICT) measurement of shock wave interaction around a circular cylinder

Three-dimensional (3-D) laser interferometric computed tomography (LICT) measurement has been applied to high-speed and unsteady flow including shock waves. The purpose of our investigation is to clarify 3-D flow phenomena quantitatively by using LICT. We used a diaphragm-less shock tube to clarify the unsteady shock wave behavior with the shot by shot method. The diaphragm-less shock tube enables us to obtain shock wave propagation reproduction in quick succession. By LICT the 3-D internal density structure of the shock–vortex flow field can be clarified. In this LICT method the observation system consists of a CCD camera, Mach–Zehnder interferometer, and nitrogen pulse laser as a light source with a suitable delay signal controller. The 3-D density CT data can be reconstructed typically from the 19 2-D density images which have a 5° interval among every shot of the diaphragm-less shock tube. As for the method to get many projection images, the optical system is fixed, while the model with a cylindrical duct and a cylinder is rotated by a fixed degree of intervals in the designated observation angle for every shot. In this paper the shock wave interaction with a finite length circular cylinder is observed by LICT measurement. For reconstruction of density distribution the Algebraic Reconstruction Technique (ART) is applied to reduce the artifact, together with precise observation.

Evolution of enstrophy in shock/homogeneous turbulence interaction

AbstractInteraction of turbulent fluctuations with a shock wave plays an important role in many high-speed flow applications. This paper studies the amplification of enstrophy, defined as mean-square fluctuating vorticity, in homogeneous isotropic turbulence passing through a normal shock. Linearized Navier–Stokes equations written in a frame of reference attached to the unsteady shock wave are used to derive transport equations for the vorticity components. These are combined to obtain an equation that describes the evolution of enstrophy across a time-averaged shock wave. A budget of the enstrophy equation computed using results from linear interaction analysis and data from direct numerical simulations identifies the dominant physical mechanisms in the flow. Production due to mean flow compression and baroclinic torques are found to be the major contributors to the enstrophy amplification. Closure approximations are proposed for the unclosed correlations in the production and baroclinic source terms. The resulting model equation is integrated to obtain the enstrophy jump across a shock for a range of upstream Mach numbers. The model predictions are compared with linear theory results for varying levels of vortical and entropic fluctuations in the upstream flow. The enstrophy model is then cast in the form of$k$–$\epsilon $equations and used to compute the interaction of homogeneous isotropic turbulence with normal shocks. The results are compared with available data from direct numerical simulations. The equations are further used to propose a model for the amplification of turbulent viscosity across a shock, which is then applied to a canonical shock–boundary layer interaction. It is shown that the current model is a significant improvement over existing models, both for homogeneous isotropic turbulence and in the case of complex high-speed flows with shock waves.

Numerical Visualization of the Unsteady Shock Wave Flow Field in Micro Shock Tube

Recently micro shock tube is extensively being used in many diverse fields of engineering applications but the detailed flow physics involved in it is hardly known due to high Knudsen number and strong compressibility effects. Unlike the macro shock tube, the surface area to volume ratio for a micro shock tube is very large. This unique effect brings many complexities into the flow physics that makes the micro shock tube different compared with the macro shock tube. In micro shock tube, the inter- molecular forces of working gas can play an important role in specifying the flow characteristics of the unsteady shock wave flow which is essentially generated in all kinds of shock tubes. In the present study, a CFD method was used to predict and visualize the unsteady shock wave flows using the unsteady compressible Navier-Stokes equations, furnished with the no-slip and slip wall boundary conditions. Maxwell's slip equations were used to mathematically model the shock movement at high Knudsen number. The present CFD results show that the propagation speed of the shock wave is directly proportional to the initial pressure and diameter of micro shock tube.

Effect of condensing moist air on unsteady shock waves around a symmetric disk in transonic internal flows

Shock waves are often observed around a symmetric disk in transonic or supersonic internal flow fields. In these cases, shock wave-boundary layer interaction becomes complex and induces the unsteady self-excited shock oscillations in the flow field. However, when water vapor is contained in the main flow, rapid expansion of the flow may give rise to non-equilibrium condensation of moist air. In the present study, the effect of condensing moist air flows on unsteady behavior of shock waves around a symmetric disk is studied numerically. The flow fields were investigated at two pressure ratios (total back pressure to total reservoir pressure) of 0.739 and 0.500. Results showed that in the absence of non-equilibrium condensation, the shock wave oscillates alternatively between the upper and lower disk surfaces at the pressure ratio of 0.739 and in the wake region at 0.500. However, in the presence of non-equilibrium condensation, the flow field aerodynamic instabilities such as root mean square of pressure oscillation and shock induced oscillations frequencies were reduced significantly compared with those in the absence of non-equilibrium condensation. In addition, the total pressure loss increased and wake frequency was reduced in case with condensing moist air flows.

Simulation of shock wave buffet and its suppression on an OAT15A supercritical airfoil by IDDES

In the present paper, extremely unsteady shock wave buffet induced by strong shock wave/boundary-layer interactions (SWBLI) on the upper surface of an OAT15A supercritical airfoil at Mach number of 0.73 and angle of attack of 3.5 degrees is first numerically simulated by IDDES, one of the most advanced RANS/LES hybrid methods. The results imply that conventional URANS methods are unable to effectively predict the buffet phenomenon on the wing surface; IDDES, which involves more flow physics, predicted buffet phenomenon. Some complex flow phenomena are predicted and demonstrated, such as periodical oscillations of shock wave in the streamwise direction, strong shear layer detached from the shock wave due to SWBLI and plenty of small scale structures broken down by the shear layer instability and in the wake. The root mean square (RMS) of fluctuating pressure coefficients and streamwise range of shock wave oscillation reasonably agree with experimental data. Then, two vortex generators (VG) both with an inclination angle of 30 degrees to the main flow directions are mounted in front of the shock wave region on the upper surface to suppress shock wave buffet. The results show that shock wave buffet can be significantly suppressed by VGs, the RMS level of pressure in the buffet region is effectively reduced, and averaged shock wave position is obviously pushed downstream, resulting in increased total lift.

Application of Hybrid ART to Laser Interferometric CT on Unsteady Shock Waves Discharged from Two Parallel and Cylindrical Nozzles

The three-dimensional (3-D) density distribution of unsteady flow field was observed by using Laser Interferometiric Computed Tomography (LICT) measurement. A diaphragmless shock tube was used to produce shock waves with high reproducibility. In our previous study, density distribution of flow field was reconstructed by ART (Algebraic Reconstruction Technique). In this paper Hybrid ART is applied to LICT measurement to reduce the artifact of reconstructed image. In our case, a computational fluid dynamics (CFD) simulation is applied to the 3-D flow fields as a CFD. Hybrid ART is the combination of CFD and ART. The reconstructed 3-D density distribution from CFD is used as the initial condition of ART. As the results reconstructed images by Hybrid ART are compared with ART, together with the precise discussion. In reconstruction results by Hybrid ART, it is confirmed influence of the density distribution interpolated initial distribution.