Characteristics of the inlet with the pressure perturbation in the ramjet engine

Flows in a ramjet inlet is simulated for the study of the rocket-ramjet transition. The flow is unsteady, two-dimensional axisymmetric, compressible and turbulent. Double time marching method is used for the unsteady calculation and HLLC method is used as a higher order MUSCL method. As for turbulent calculation, k — ω SST model is used for more accurate viscous calculations. Sinusoidal pressure perturbation is given at the exit and the flow fields at the inlet is studied. The cruise condition as well as the ground test condition are considered. The pressure level for the ground test condition is relatively low and the effect of the pressure perturbation at the combustion chamber is small. The normal shock at the cruise condition is very sensitive to the pressure perturbation and can be easily detached from the cowl when the exit pressure is relatively high. The sudden decrease in the mass flux is observed when the inlet flow becomes subcriticai, which can make the inlet incapable. The amplitude of travelling pressure waves becomes larger as the downstream pressure increases, and the wavelength becomes shorter as Mach number increases. The phase difference of the travelling perturbed pressure wave in space is 180 degree.     

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.     

Experimental study on hysteresis phenomena of shock wave structure in an over-expanded axisymmetric jet

In recent studies on two-dimensional supersonic jets, it is reported that the hysteresis phenomenon for the reflection type of shock waves in the jet flow field is occurred under the quasi-steady flow condition and this phenomenon is affected by the transitional pressure ratio between the regular reflection and Mach reflection. However, so far, there are few researches on the hysteresis phenomenon for the transition of shock waves between regular and Mach reflection in over-expanded supersonic jets and the phenomenon has not been investigated satisfactorily. Therefore, the purpose of this study is to clarify the hysteresis phenomena for the reflection type of shock wave in the over-expanded axi-symmetric supersonic jet experimentally, and to discuss the relationship between hysteresis phenomenon and rate of the change of pressure ratio with time. Furthermore, the effect of Mach number at the nozzle exit on hysteresis loop was investigated for two kinds of nozzle.     

Numerical investigation on the directionality of nonlinear indicial responses

An unsteady Euler solver is modified to investigate the directionality of nonlinear indicial response to a step change in the angle of attack. An impulsive change in the angle of attack is incorporated by using the field velocity approach, which is known to decouple the step change in the angle of attack from a pitch rate. Numerical results are thoroughly compared against analytical results for two-dimensional indicial responses. The same method is applied to investigate the directionality of nonlinear indicial responses. It is found that directionality is mainly due to the asymmetry of initial shock locations. Since the directionality of the pitching moment responses is significant in the critical Mach number region, it is also shown that consideration of the directionality is crucial for accurate modeling of the nonlinear indicial functions.     

The effect of the secondary annular stream on supersonic jet

The present study addresses an experimental investigation of the near field flow structures of supersonic, dual, coaxial, free, jet, which is discharged from the coaxial annular nozzle. The secondary stream is made from the annular nozzle of a design Mach number of 1.0 and the primary inner stream from a convergent-divergent nozzle. The objective of the present study is to investigate the interactions between the secondary stream and inner supersonic jets. The resulting flow fields are quantified by pitot impact and static pressure measurements and are visualized by using a shadowgraph optical method. The pressure ratios of the primary jet are varied to obtain over-expanded flows and moderately under-expanded flows at the exit of the coaxial nozzle. The pressure ratio of the secondary annular stream is varied between 1.0 and 4.0. The results show that the secondary annular stream significantly changes the Mach disc diameter and location, and the impact pressure distributions. The effects of the secondary annular stream on the primary supersonic jet flow are strongly dependent on whether the primary jet is underexpanded or over-expanded at the exit of the coaxial nozzle.     

An analysis of flow and screech tone from supersonic axisymmetric jet at the initial stage

The screech tone from jet Mach number 1.18 is numerically calculated from the initial stage. A fourth-order optimized compact scheme and fourth-order Runge-Kutta method are used to solve the 2D axisymmetric Euler equation. Pulse jet problem with Mach number 1.56 is solved to validate the present method. Not only two important components of generating screech tones, shock cell structure and vortices, but also the transient behavior of the screech tone are investigated at the initial stage. Additional time is necessary to generate periodic screech tone after formation of shock cell structures. As time goes on, the location of the vortex generation becomes fixed near the nozzle exit, which is farther downstream initially, and the screech tone becomes periodic. The FFT results of three different periods are compared. It is observed that the component of lower frequency is dominant at the beginning, and the component of the screech tone becomes dominant as time increases. It can be concluded that the screech tones can be also numerically reproduced at the initial stage with the present inviscid method.     

Transitional behavior of a supersonic flow in a two-dimensional diffuser

Two-dimensional blow-down type supersonic wind tunnel was designed and built to investigate the transient behavior of the startup of a supersonic flow from rest. The contour of the divergent part of the nozzle was determined by the MOC calculation. The converging part of the nozzle, upstream of the throat was contoured to make the flow profile uniform at the throat. The flow characteristics of the steady supersonic condition were visualized using the highspeed schlieren photography. The Mach number was evaluated from the oblique shock wave angle on a sharp wedge with half angle of 5 degree. The measured Mach number was 2.4 and was slightly less than the value predicted by the design calculation. The initial transient behavior of the nozzle was recorded by a high-speed digital video camera with schlieren technique. The measured transition time from standstill to a steady supersonic flow was estimated by analyzing the serial images. Typical transition time was approximately O.1sec.     

The self-induced oscillations of the under expanded jets impinging upon a cylindrical body

The present study addresses the flow characteristics involved in the self-induced oscillations of the underexpanded jet impinging upon a cylindrical body. Both experiment and computational analysis are carried out to elucidate the shock motions of the self-induced oscillations and to find the associated major flow factors. The underexpanded sonic jet is made from a nozzle and a cylindrical body is placed downstream to simulate the impinging jet upon an obstacle. The computational analysis using TVD scheme is applied to solve the axisymmetric, unsteady, inviscid governing equations. A Schlieren system is employed to visualize the self-induced oscillations generated in flow field. The data of the shock motions are obtained from a high-speed video system. The detailed characteristics of the Mach disk oscillations and the resulting pressure variations are expatiated using the time dependent data of the Mach disk positions. The mechanisms of the self-induced oscillations are discussed in details based upon the experimental and computational results.     

Numerical analysis of impulse waves discharged at the exit of a model tunnel

Impulse waves are micro-pressure waves, which always occur at the tunnel exit when a high-speed train is moving inside a train tunnel. The air around the train nose is compressed and compression waves are induced. The impulse wave is discharged at the exit of a train tunnel when a compression wave propagates outside of the tunnel exit. Impulse waves are weak-strength pressure waves, which lead to noise and other environmental problems. In order to efficiently control the impulse wave at the exit of a train tunnel, numerical studies on investigating the generation and propagation of the impulse wave were carried out. A 2-D axisymmetric model tunnel was simulated at different operating conditions. Different Mach numbers of compression waves were varied to induce different magnitudes of impulse waves at the tunnel exit. In addition, compression waves with different pressure gradients were assumed at the tunnel entry to check their effects on the generation of impulse waves. In order to observe impulse waves at far field, five monitor points were installed behind the tunnel exit to record pressure histories as impulse waves moved through these locations. The detailed magnitudes and characteristics of impulse waves were obtained in the present studies.     

Study of moist air flow through the ludwieg tube

The time-dependent behavior of unsteady condensation of moist air through the Ludwieg tube is investigated by using a computational fluid dynamics (CFD) work. The twodimensional, compressible, Navier-Stokes equations, fully coupled with the condensate droplet growth equations, are numerically solved by a third-order MUSCL type TVD finite-difference scheme, with a second-order fractional time step. Baldwin-Lomax turbulence model is employed to close the governing equations. The predicted results are compared with the previous experiments using the Ludwieg tube with a diaphragm downstream. The present computations represent the experimental flows well. The time-dependent unsteady condensation charac-teristics are discussed based upon the present predicted results. The results obtained clearly show that for an initial relative humidity below 30% there is no periodic oscillation of the condensation shock wave, but for an initial relative humidity over 40% the periodic excursions of the condensation shock occurs in the Ludwieg tube, and the frequency increases with the initial relative humidity. It is also found that total pressure loss due to unsteady condensation in the Ludwieg tube should not be ignored even for a very low initial relative humidity and it results from the periodic excursions of the condensation shock wave.     

Passive prandtl-meyer expansion flow with homogeneous condensation

Prandtl-Meyer expansion flow with homogeneous condensation is investigated experimentally and by numerical computations. The steady and unsteady periodic behaviors of the diabatic shock wave due to the latent heat released by condensation are considered with a view of technical application to the condensing flow through steam turbine blade passages. A passive control method using a porous wall and cavity underneath is applied to control the diabatic shock wave. Two-dimensional, compressible Navier-Stokes with the nucleation rate equation are numerically solved using a third-order TVD (Total Variation Diminishing) finite difference scheme. The computational results reproduce the measured static pressure distributions in passive and no passive Prandtl-Meyer expansion flows with condensation. From both the experimental and computational results, it is found that the magnitude of steady diabatic shock wave can be considerably reduced by the present passive control method. For no passive control, it is found that the diabatic shock wave due to the heat released by condensation oscillates periodically with a frequency of 2.40 kHz. This unsteady periodic motion of the diabatic shock wave can be completely suppressed using the present passive control method.     

Computational study of the unsteady flow characteristics of a micro shock tube

Micro shock tubes are widely employed in many micro instruments which require high speed and high temperature flow field. The small flow dimension introduces additional flow physics such as rarefaction effects, viscous effects etc, which makes the micro shock tube different from conventional macro shock tubes. In the present study, a numerical investigation of the flow physics associated with shock propagation and reflection inside micro shock tubes was carried out using unsteady Navier Stokes equations. Maxwell’s slip boundary conditions were incorporated to simulate the rarefaction effects produced due to low pressure and very small length scale. The effect of initial pressures on the shock propagation was investigated keeping the pressure ratio constant. The dependency of the shock tube diameter on shock propagation was also investigated. The results show that shock strength attenuates drastically in a micro shock tube compared to macro shock tubes. The viscous boundary layer becomes a governing parameter in controlling micro shock tube wave propagations. The implementation of slip velocity to model rarefaction effects increases the shock strength and aids in shock wave propagation. The simulation with slip wall exhibits a wider hot zone (shock-contact distance) compared to no-slip simulation. The contact surface propagation distance reduces under the slip effects. A drastic attenuation in shock propagation distance was observed with reduction in diameter. The shock wave when reflected from the end wall inhibits the rarefaction effects, generally happening at very low pressure micro shock tubes, and the associated slip effect vanishes for the post reflected shock flow field.     

Effect of boundary layer swirl on supersonic jet instabilities and thrust

This paper reports the effects of nozzle exit boundary layer swirl on the instability modes of underexpanded supersonic jets emerging from plane rectangular nozzles. The effects of boundary layer swirl at the nozzle exit on thrust and mixing of supersonic rectangular jets are also considered. The previous study was performed with a 30° boundary layer swirl (S=0.41) in a plane rectangular nozzle exit. At this study, a 45° boundary layer swirl (S=1.0) is applied in a plane rectangular nozzle exit. A three-dimensional unsteady compressible Reynolds-Averaged Navier-Stokes code with Baldwin-Lomax and Chien’sk-ε two-equation turbulence models was used for numerical simulation. A shock adaptive grid system was applied to enhance shock resolution. The nozzle aspect ratio used in this study was 5.0, and the fully-expanded jet Mach number was 1.526. The “flapping” and “pumping” oscillations were observed in the jet’s small dimension at frequencies of about 3,900Hz and 7,800Hz, respectively. In the jefs large dimension, “spanwise” oscillations at the same frequency as the small dimension’s “flapping“ oscillations were captured. As reported before with a 30° nozzle exit boundary layer swirl, the induction of 45° swirl to the nozzle exit boundary layer also strongly enhances jet mixing with the reduction of thrust by 10%.     

Characteristic analysis of supersonic impinging jet in laser machining

The interaction of a gas jet with a workpiece in laser machining is investigated by studying the influence of the processing parameters on the dynamic characteristic of the gas flow in the hole, mainly including the mass flow rate and the thrust. The modeling of a supersonic turbulent jet impinging on a plate with a hole concentric with the jet is presented. Numerical simulations are carried out using an explicit, coupled solution algorithm with solution-based mesh adaptation. The model is able to make quantitative predictions of the effect of the standoff distance on the mass flow rate and the axial thrust. It is revealed that the shape of the hole has weak effects on shock structure, but it can improve the dynamic characteristic slightly. The standoff distance has great effect not only on the shock structure, but also on the jump and fluctuation of the thrust for the different exit Mach number.     

A study on impulsive sound attenuation for a high-pressure blast flowfield

The present work addresses a numerical study on impulsive sound attenuation for a complex high-pressure blast flowfield; these characteristics are generated by a supersonic propellant gas flow through a shock tube into an ambient environment. A numerical solver for analyzing the high pressure blast flowfield is developed in this study. From numerical simulations, wave dynamic processes (which include a first precursor shock wave, a second main propellant shock wave, and interactions in the muzzle blasts) are simulated and discussed. The pressure variation of the blast flowfield is analyzed to evaluate the effect of a silencer. A live firing test is also performed to evaluate four different silencers. The results of this study will be helpful in understanding blast wave and in designing silencers.     

Influences of mach number and flow incidence on aerodynamic losses of steam turbine blade

An experiment was conducted to investigate the aerodynamic losses of high pressure steam turbine nozzle (526A) subjected to a large range of incident angles (−34° to 26°) and exit Mach numbers (0.6 and 1.15). Measurements included downstream Pitot probe traverses, upstream total pressure, and endwall static pressures. Flow visualization techniques such as shadowgraph and color oil flow visualization were performed to complement the measured data. When the exit Mach number for nozzles increased from 0.9 to 1.1 the total pressure loss coefficient increased by a factor of 7 as compared to the total pressure losses measured at subsonic conditions (M₂<0.9). For the range of incidence tested, the effect of flow incidence on the total pressure losses is less pronounced. Based on the shadowgraphs taken during the experiment, it’s believed that the large increase in losses at transonic conditions is due to strong shock/ boundary layer interaction that may lead to flow separation on the blade suction surface.     

An experimental study of supersonic dual coaxial free jet

A supersonic dual coaxial jet has been employed popularly for various industrial purposes, such as gasdynamic laser, supersonic ejector, noise control and enhancement of mixing. Detailed characteristics of supersonic dual coaxial jets issuing from an inner supersonic nozzle and outer sonic nozzles with various ejection angles are experimentally investigated. Three important parameters, such as pressure ratios of the inner and outer nozzles, and outer nozzle ejection angle, are chosen for a better understanding of jet structures in the present study. The results obtained from the present experimental study show that the Mach disk diameter becomes smaller, and the Mach disk moves toward the nozzle exit, and the length of the first shock cell decreases with the pressure ratio of the outer nozzle. It was also found that the highly underexpanded outer jet produces a new oblique shock wave, which makes jet structure much more complicated. On the other hand the outer jet ejection angle affects the structure of the inner jet structure less than the pressure ratio of the outer nozzle, relatively.     

A numerical study of shock wave/boundary layer interaction in a supersonic compressor cascade

A numerical analysis of shock wave/boundary layer interaction in transonic/supersonic axial flow compressor cascade has been performed by using a characteristic upwind Navier-Stokes method with various turbulence models. Two equation turbulence models were applied to transonic/supersonic flows over a NACA 0012 airfoil. The results are superion to those from an algebraic turbulence model. High order TVD schemes predicted shock wave/boundary layer interactions reasonably well. However, the prediction of SWBLI depends more on turbulence models than high order schemes. In a supersonic axial flow cascade at M=1.59 and exit/inlet static pressure ratio of 2.21, k-μ and Shear Stress Transport (SST) models were numerically stables. However, the k-μ model predicted thicker shock waves in the flow passage. Losses due to shock/shock and shock/boundary layer interactions in transonic/supersonic compressor flowfields can be higher losses than viscous losses due to flow separation and viscous dissipation.     

A high resolution numerical scheme for a high speed gas-liquid two-phase flow

A high resolution numerical method for solving high speed gas-liquid two-phase flow is proposed and applied to the two-phase shock tube problem. The present method employs a finite-difference 4th-order Runge-Kutta method and Roe’s flux difference splitting approximation with the MUSCL TVD scheme. A homogeneous equilibrium gas-liquid two-phase model that takes account of the compressibility of mixed media is used. Therefore, the present density-based numerical method permits simple treatment of the whole gasliquid two-phase flow field, including wave propagation, large density changes and incompressible flow characteristics at the low Mach number. The speed of sound of above two-phase media has been derived on the basis of thermodynamic relations. By this method, a Riemann problem for the Euler equations of a one-dimensional shock tube was computed. Numerical results such as detailed observations of shock and expansion wave propagations through the gas-liquid two-phase media at thermal and isothermal conditions, and some features related to computational efficiency are made. Comparisons of predicted results with exact solutions are provided and discussed.     

静压气体轴承中的激波与边界层相互影响的研究进展

综述空气动力学理论中激波—边界层相互影响的一般特性以及静压气体轴承中的激波与边界层相互影响理论与试验研究。指出随着计算技术和微流体测试技术的发展,气膜内的压力分布、速度分布和马赫数分布,根据流场各段不同的流动特点,将可以采用合适的湍流模型进行计算机模拟,并采用相关试验得到验证。静压气体轴承中的激波—边界层相互影响将得到更深入的研究,必将促进对较高供气压力和较大气膜间隙条件下,静压气体轴承气膜入口、出口转角区和气膜内轴承特性的全面了解。