Correlations study in shock wave/turbulent boundary layer interaction

The shock waves/turbulent boundary layer interaction is a problem of critical importance that is frequently encountered in designing flying vehicles. Presently, the most topical issue is the investigation of nonstationary phenomena (in particular, low-frequency effects) involved in this interaction. We have experimentally studied separated flows in the zone of interaction between an obliquely incident shock wave and a turbulent boundary layer at a Mach number of M = 2. Correlation data in the separation zone and the upstream flow were obtained. It is established that low-frequency oscillations in the reflected shock are related to the pulsation in the inflow boundary layer.     

Supersonic Flow of Viscous Gas in a Flat Channel at High Values of the Reynolds Number

Based on numerical analysis of two-dimensional Reynolds equations using the differential model of turbulence, the structure of flow field and aerodynamic characteristics of a flat channel with variable cross section at the inlet Mach number of four is investigated in the range of Reynolds numbers Re = 10⁵–10⁷. According to the calculation results, the interaction between a shock and a laminar boundary layer results in the emergence of a closed separation zone. During interaction with a turbulent boundary layer, two flow schemes are possible, depending on the intensity of incident shock, namely, without and with separation. The extrapolation of calculation data to nonviscous flow (limiting transition Re ) shows that the classical scheme of regular reflection of the shock from a flat surface corresponds to interaction without separation. Corresponding to interaction with separation is the flow scheme with formation of a small closed separation zone in which a subsonic circulation flow takes place.     

The interaction of shock waves with a boundary layer on a sharp plate and a blunted plate

The interaction of shock waves with a turbulent boundary layer on a sharp plate and a blunted plate is numerically investigated. The shock waves in the flow are generated by wedges installed on the flat plate. The flow is simulated by the dynamic equations of a viscous perfect gas. The effect of the blunting radius of the plate’s leading edge and the wedge angle on the flow field and the local aerodynamic coefficients is shown. The calculated results are in agreement with the experimental data.     

An experimental investigation of the Mach number effect upon a normal shock wave — Tubulent boundary layer interaction on a curved wall

Summary The Mach number effect upon the normal shock wave — turbulent boundary interaction is investigated experimentally.The measurements included flow parameters distribution determination, boundary layer development through the interaction area and the shock wave topography visualization. In order to have better insight into separation development the oil flow technique has been applied. The comparison of our results with other published data is presented.With 10 FiguresThe experiments have been carried out at the Institut für Strömungslehre und Strömungsmaschinen, Universität Karlsruhe, Federal Republic of Germany     

An experimental investigation of the Reynolds number effect on a normal shock wave-turbulent boundary layer interaction on a curved wall

Summary The presented experiment concerned an interaction between a normal shock wave, terminating a local supersonic area in a curved duct, and a turbulent boundary layer developed along the convex wall. This paper deals with the Reynolds number effect upon the interaction structure.The measurements included flow parameters distribution determination, boundary layer development through the interaction area and the shock wave topography visualization. In order to gain more information about separation the wall oil tracing has been applied. The comparison of our results with other published data is presented.With 11 FiguresThe experiments have been carried out at the Institut für Strömungslehre und Strömungsmaschinen, Karlsruhe Universität, Karlsruhe, Federal Republic of Germany.     

Sliding electric arc discharge as a means of aircraft trajectory control

The dynamics of sliding electric arc discharge and the formation of shock waves in the stages of leader motion and the electric arc development in a supersonic air flow behind the shock wave have been studied for an initial pressure of 0.09–0.5 atm (bar). The air flow in the discharge was imaged using an optical system comprising a shadow device (IAB-458), an optical interference attachment (RP-452), and a modified ruby laser (OGM-20) producing 10–15 output pulses per pumping pulse. Stable initiation of sliding electric arc discharge takes place in a supersonic air flow behind the shock waves with 1.7<M<3.4. This discharge produces shock waves leading to separation of the boundary layer and to an increase in the pressure at the surface. These shock waves can be used for modifying gasdynamics in the air flow streamlining the surface and for controlling the motion of an aircraft.     

Interaction of a shock wave with a turbulent boundary layer on a heated surface

The interaction of a shock wave with a turbulent boundary layer has been studied. A stream has been diagnosed using non-invasive visualization (schlieren method). During the investigation the effect of surface heating on the length of boundary layer separation has been studied. It has been discovered that during surface heating the dependence of the length of separation on the temperature ratio is approximated using a linear law.     

Laminar-turbulent transition in a supersonic boundary layer during initiation of a pulsed surface discharge

The near-surface layer structure in a supersonic airflow (M = 1.5) behind the plane shock wave has been experimentally studied in a shock tube. The flow structure was visualized using nanosecond pulsed distributed surface discharge. Different structures of plasma glow have been observed for the laminar and turbulent flow regimes in the boundary layer. The position of a region of the laminar-turbulent transition for different flow densities (0.11 and 0.19 kg/m³) has been determined and the critical Reynolds number (Re _k ～ 2.4 × 10⁵) for this transition has been evaluated.     

某跨音速高压涡轮非定常流动数值模拟研究

通过求解三维非定常雷诺平均N-S方程模拟某跨音速高压涡轮非定常流场,研究涡轮内非定常流动特征。通过对静子尾迹及静子尾缘激波和转子叶排之间的相互干涉过程进行详细分析,发现定常/非定常模拟方法获得的涡轮总体性能参数基本一致但流场存在较大差异。静子尾迹是导致涡轮流场非定常性的重要因素之一:在转子叶栅通道中部和下部,静子尾迹和转子叶片附面层及下通道涡发生明显干涉,并导致通道中下部损失周期性波动幅度较大,此外尾迹和下通道涡间的干涉作用在转子尾缘处诱导出高频脱落涡。静子尾缘激波也是导致涡轮流场非定常性的原因之一,激波和转子叶片作用形成复杂的波系结构,对涡轮流场影响显著:一方面激波/附面层干涉导致转子和静子的吸力面产生周期性变化的高温区域;另一方面激波撞击叶片导致叶片表面的气流在激波后出现分离,对转子静压分布产生影响,使得转子叶片表面载荷出现明显的非定常性,进而导致涡轮输出功的周期性波动十分剧烈。     

The structure of a turbulent boundary layer

According to the wave mechanism of turbulence, pulsation in the hydrodynamic parameters results from a superposition of perturbations arising at the wall and then spreading in the flow in the form of spherical wave packets. At the flow boundary, where the fluid velocity is characterized by a large gradient, the acoustic rays of these waves exhibit bending and reversal toward the wall, whereby the trajectories with various initial orientations are interweave and the wave packets are broken. The pulsation of parameters in the region of wave packet breakage results in the formation of a turbulent boundary layer. Upon the reflection of waves, the flow velocity oscillations immediately at the wall cease that corresponds to a laminar sublayer of the turbulent boundary layer.     

An efficient numerical method for highly loaded transonic cascade flow

In this paper, an efficient numerical method for transonic viscous flow in a highly loaded turbine vane cascade, where the interaction of a shock wave and boundary layer often leads to very complicated flow phenomena, is developed. The numerical code, a modified implicit flux-vector-splitting solver of the Navier–Stokes equations (MIFVS), is extended to simulate such transonic cascade flow. A compressible low-Reynolds-number k–ɛ model, together with a transition-modified damping function, has been implemented into the MIFVS code. With this extended MIFVS solver, the main feature of transonic flow and shock and boundary-layer interactions in the highly loaded transonic turbine vane are efficiently predicted with satisfactory accuracy. The convergence rate is found to be three times faster than that of flux-vector-splitting (FVS) methods.     

湍流边界层激励下加筋平板振声响应特性研究

船舶、汽车和飞机等高速运动时,其外壳受湍流边界层壁面脉动压力激励而产生的内场声辐射成为该类交通工具自噪声的重要成分。基于模态叠加法计算结构振动响应。采用湍流壁面脉动压力功率谱Corcos模型,计算了外侧气流或水流湍流边界层激励下简支平板振动及内场辐射声,计算值与解析解和试验值吻合良好,验证了算法的有效性。采用湍流壁面脉动压力功率谱改进型Corcos模型,研究了外侧水流湍流边界层激励下平板及板格的振声响应特性,结果表明:水流马赫数低,壁面脉动压力迁移波数大于平板结构弯曲波数,壁面脉动压力波数-频率谱的迁移脊对平板的激励作用可以忽略;横向或纵向加筋对板格振动速度自功率谱级基本无影响;减小板格宽度与长度之比,适当增大板格流向长度可使平板振动辐射声功率在2 000 Hz以上明显降低。     

高超声速激波/边界层干扰及MVG阵列流动控制研究

高超声速飞行器流场中通常会伴随激波/边界层干扰(SWBLI),其引发的流动分离将导致进气道性能下降。该文采用湍流分离涡(DES)方法、结合有限体积离散方法对来流马赫数为7流场中SWBLI诱导的分离气泡进行数值研究,模拟结果清晰地显示了分离气泡从产生到充分发展的具体过程,揭示了分离气泡的产生机理。利用微型涡流发生器(MVG)阵列对其进行控制,讨论了流场结构、壁面静压力、壁面剪切力及总压损失等参数变化对SWBLI控制效果的影响。结果表明:MVG阵列可显著改变高超声速流体边界层,使得分离气泡尺寸减小,分离激波强度减弱,分离气泡内及其下游流体的流向速度梯度增加,总压损失降低可达1.9%。     

Edge effects in flow around a plasma actuator

We have studied the structure of flow formed in the boundary layer at the lateral edge of the discharge zone of a dielectric barrier discharge plasma actuator. It is established that a region with nonzero component of tangential velocity exists near the plasma layer boundary. At some distance downstream of the actuator, a concentrated vortex is formed with the axis aligned with the flow direction. In the presence of two closely spaced boundaries of the plasma region, a pair of counter-rotating vortices is formed. Separate microdischarges in the plasma layer also appear to be a source of similar longitudinal vortices with smaller amplitude.

     

Wind loads on a residential solar water heater

This experimental study aims to address the effects of freestream turbulence intensity (TI) on wind loads on a residential solar water heater with and without a guide plate. Highly turbulent flow was produced by a turbulence generation grid at the inlet of the working section. Results show that the distributions of mean longitudinal and spanwise pressures are associated with tilt angle, guide plate, and freestream TI. Interactions between separation and corner vortices on the upper surface and impingement of shear layer from the tip of a guide plate on the lower surface would induce higher levels of fluctuating pressure, particularly in turbulent flow. A similarity parameter α/A* is also proposed to scale the uplift coefficients in both smooth uniform and turbulent flows.     

Three-dimensional aspects of boundary-layer transition

A review is made in this paper of the three-dimensional nature of instability leading to transition in a two-dimensional flow on a flat plate (Blasius boundary layer) or between parallel plates (plane Poiseuille flow), with additional reference to the flow on a concave wall. Instability and transition in a three-dimensional boundary layer are then reviewed, with particular attention to the flow due to a rotating disk and the flow perturbed by a three-dimensional roughness element on a flat plate. The growth of a turbulent spot is discussed as a phenomenon exhibiting a similar feature of instability. Only the flow of an incompressible fluid is considered.     

Nonstationary phenomena in the region of shock-wave interaction with a boundary layer at transonic flow velocities

Nonstationary characteristics of detached flow have been experimentally studied during interaction of the boundary layer with a shock wave that appears on a profiled bump in transonic flow. The experiments were performed with variable shock-wave intensity and position in a T-325 wind tunnel. The flow was studied using methods of schlieren imaging, measuring average pressure and its pulsations on the surface of a model, and determining velocity fields by particle image velocimetry. Analysis of the experimental data showed that the observed shock-wave oscillations and flow pulsations in the detachment zone were related to disturbances present in the oncoming boundary layer.     

Investigating effect of temperature and time of metalised layer sintering by SMPP method on tensile strength and thermal shock resistance

Abstract

Sintered metal powder process is one of the high technology methods in ceramic–metal joining processes. Improvement in joining zone properties is very important in this method. The present study reveals the effect of metalised layer sintering temperature and time, and applied layer thickness on tensile strength and thermal shock resistance of alumina–copper joint. The results reveal that primary sintering for holding time duration of 90 min at a temperature of 1530°C and applied layer thickness of 50 μm with proper different stages of plating and brazing leads to a tensile strength of 120 MPa in the joining zone. The specimens, which were joined in this condition, were thermal shock resistance.     

Simulation of Turbulent Flow Past Bluff Bodies on Coarse Meshes Using General Galerkin Methods: Drag Crisis and Turbulent Euler Solutions

In recent years adaptive stabilized finite element methods, here referred to as General Galerkin (G2) methods, have been developed as a general methodology for the computation of mean value output in turbulent flow. In earlier work, in the setting of bluff body flow, the use of no slip boundary conditions has been shown to accurately capture the separation from a laminar boundary layer, in a number of benchmark problems. In this paper we extend the G2 method to problems with turbulent boundary layers, by including a simple wall-model in the form of a friction boundary condition, to account for the skin friction of the unresolved turbulent boundary layer. In particular, we use G2 to simulate drag crisis for a circular cylinder, by adjusting the friction parameter to match experimental results. By letting the Reynolds number go to infinity and the skin friction go to zero, we get a G2 method for the Euler equations with slip boundary conditions, which we here refer to as an EG2 method. The only parameter in the EG2 method is the discretization parameter, and we present computational results indicating that EG2 may be used to model very high Reynolds numbers flow, such as geophysical flow.     

What makes the difference between shock boundary-layer interaction on an airfoil and on a swept wing?

Summary Swept wing flows are characterized by the curvature of the streamlines in the projection to the wing plan and by the skewing of the velocity profile in the boundary layer. The aerodynamic performance of supercritical wings at transonic speeds is trongly influenced by the interaction between a weak shock front and a turbulent boundary layer. The characteristic elements of this interaction are the precompression, the post-shock expansion, and the shock diffusion. The differences between the interactive flow over an airfoil and over a swept wing are elaborated by the comparison between the two-dimensional case and the flow with superposed tangential velocity.