前体涡发生器对轴对称高超声速进气道激波振荡流动的影响实验

激波振荡是高超声速进气道不起动过程中常见的流动现象,会显著降低进气道气流捕获与压缩效率、产生剧烈的非定常气动力载荷而危害飞行器安全. 从激波振荡的控制出发,实验研究了前体转捩带位置的涡发生器对轴对称高超声速进气道激波振荡流动的影响. 分别在起动和激波振荡两种进气道流态下,选择无、0.5 mm与1 mm高度涡发生器工况进行对比研究. 并采用高速纹影与壁面动态测压同步记录非定常流动特征. 结果表明,1 mm高度内的涡发生器对起动状态的进气道主流流场结构、壁面压强分布影响不显著. 但对于激波振荡流动,涡发生器会明显缩小外压缩面分离区运动范围,缩短振荡周期,提升振荡周期内壁面压强的时均值. 涡发生器的影响程度随其高度的增大而增强,其中振荡周期从无涡发生器的4 ms缩短到1 mm高度涡发生器的3.13 ms. 此外,0.5 mm高度涡发生器会使得进气道内部测点的压强振荡幅值整体下降,相比无涡发生器工况的下降幅度可达23%. 流场结构与壁面压强信号的分析表明,涡流发生器主要通过其产生的流向涡影响激波振荡流动,包含流向涡对下游边界层的扰动以及流向涡与分离区的相互干扰.      

前体涡发生器对轴对称高超声速进气道激波振荡流动的影响实验

激波振荡是高超声速进气道不起动过程中常见的流动现象,会显著降低进气道气流捕获与压缩效率、产生剧烈的非定常气动力载荷而危害飞行器安全.从激波振荡的控制出发,实验研究了前体转捩带位置的涡发生器对轴对称高超声速进气道激波振荡流动的影响.分别在起动和激波振荡两种进气道流态下,选择无、0.5 mm与1 mm高度涡发生器工况进行对比研究.并采用高速纹影与壁面动态测压同步记录非定常流动特征.结果表明,1 mm高度内的涡发生器对起动状态的进气道主流流场结构、壁面压强分布影响不显著.但对于激波振荡流动,涡发生器会明显缩小外压缩面分离区运动范围,缩短振荡周期,提升振荡周期内壁面压强的时均值.涡发生器的影响程度随其高度的增大而增强,其中振荡周期从无涡发生器的4 ms缩短到1 mm高度涡发生器的3.13 ms.此外,0.5 mm高度涡发生器会使得进气道内部测点的压强振荡幅值整体下降,相比无涡发生器工况的下降幅度可达23%.流场结构与壁面压强信号的分析表明,涡流发生器主要通过其产生的流向涡影响激波振荡流动,包含流向涡对下游边界层的扰动以及流向涡与分离区的相互干扰.     

爆轰波通过扩张喷管的双曝光全息实验和数值研究

结合实验和数值模拟方法,对以脉冲爆轰发动机为背景的爆轰波通过扩张喷管的流动进行了系列研究。实验采用双曝光全息干涉方法对爆轰波绕射流场进行测量,得到了比传统的纹影法更清晰和可定量化的照片。发展了基于非结构四边形网格自适应有限体积程序,结合基元化学反应模型对扩张喷管中爆轰化学反应流场进行了数值模拟,模拟结果与实验照片吻合较好。实验和数值模拟结果表明,爆轰波绕射具有许多和激波绕射不同的流场特征,其中包括二次起爆现象、化学反应面与前导激波相脱离而引起的复杂流场等,同时初始压力和扩张角度变化也对爆轰波绕射过程产生较大影响,初始压力越低,化学反应区和前导激波分离现象越明显,且前导激波的曲率越大。     

侧压式道气道流场特性研究

应用表面油流和压力测量技术,结合数值模拟方法研究了超燃冲压发动机侧压式进气道无支板和有支板两种典型工况的基本流动特性.结果表明:由侧壁诱发的后掠激波与底板边界层相互作用,形成一对逆向旋转的漩涡,该漩涡在下游不断发展并偏离底板,在出口截面形成一个低马赫数低总压区;支板的引入增加了压缩率,但同时导致底板与侧壁边界层严重分离,使进气道性能偏离设计工况.     

纳米示踪平面激光散射技术在激波复杂流场测量中的应用

在激波以及激波边界层相互作用这类含激波的复杂流场中,流场结构具有明显的三维特征.研究这类流场,采用纹影、阴影和干涉等传统流动显示技术空间分辨率较低,难以分辨流场的三维特性.基于纳米示踪的平面激光散射技术(nano-tracer planar laser scattering,NPLS),是作者近年来开发的一种新的研究超声速流场的测试与显示技术,可对超声速复杂三维流场进行高时空分辨率流动显示与测量.NPLS技术的特点使其成为测量激波复杂流场的有力手段.近年来,作者以NPLS技术为主要手段,对航空航天领域典型的激波复杂流场进行了试验研究,包括超声速弹头绕流、超声速混合层、超声速边界层,以及激波边界层相互作用流场,显示出NPLS技术在激波复杂流场精细测试与流动显示中优势.本文简要介绍NPLS技术在激波复杂流场测量中应用的研究进展.      

超声速多喷流干扰流场特性研究

研究了旋成体上超声速来流与超声速横向多喷流相撞产生的层流干扰流场特性. 数 值方法针对三维可压缩Navier-Stokes方程按二阶精度Roe格式进行离散,采用基于多区对 接网格技术的有限体积法. 数值模拟结果描述了多喷流干扰流场的空间结构以及激波/边界层 干扰引起的分离范围,探讨了沿流向等间距排列的喷口个数对表面和空间流场结构以及压力 分布的影响规律. 结果表明,第一喷口对多喷流干扰流场主要结构和喷口上游表面分离范围 起主导作用. 其中三喷流流场数值模拟的对称面激波结构与实验纹影结果进行对比,符合较 好.     

管道凸起激波/湍流边界层干扰的数值模拟

采用基于部分平均的湍流方程组,模拟了管道凸起流动中两种不同强度的激波-湍流边界层干扰现象,对流场的时均参数与实验值进行了比较,计算得到的壁面压力分布、摩阻系数分布、边界层厚度和速度型与实验值比较吻合很好.结果表明基于部分平均的湍流方程组既能够准确模拟小分离的激波/湍流边界层干扰流动又能够准确模拟大分离的激波/湍流边界层干扰流动,较好地预测了典型的λ激波结构.     

抽吸激波管的性能分析与计算

为了减小由于反射激波和透射激波分叉引起的反射型激波风洞试验气体提前受到污染的现象,本文研究了一种新的具有抽吸的激波管流动,分析了抽吸缝的作用,给出了这种抽吸激波管性能参数的计算方法,同时还给出了反射激波与边界层相互作用引起的激波分叉的形状随抽吸量变化的计算公式。实验证实了边界层抽吸可以有效地减小激波与壁面边界层相互作用所产生的分离现象。计算与测量结果是一致的。     

后掠激波诱导分离特性的马赫数影响

利用壁面压力测量和表面油流显示技术对后掠压缩角产生的激波与湍流边界层干扰进行了实验研究。主要研究了来流马赫数对干扰分离特性和起始分离条件的影响，实验的马赫数为１．７９，２．０４和２．５０，相应的雷诺数为２．４２—２．４７×１０￣７／ｍ．实验模型共１５个，其后掠角变化范围是０°—６０°，流向压缩角变化范围是１０°—３０°，实验结果表明，与二维干扰不同，对于大后掠角的三维干扰来说，分离线上的压比和起始分离时的无粘激波压比近似相等，且是一个与马赫数无关的常数。     

气相爆轰波绕射流场显示研究

采用基于红宝石激光器（波长694.3 nm）的纹影系统,对气相爆轰波绕射进行了初步的流场显示研究。采用单色激光和合适半带宽（15 nm）的滤光片,有效地消除了爆轰波自发光对流场显示的影响。合理设置激光器同步控制系统的触发延时,得到了序列的爆轰波阵面纹影照片。结果表明:图像清晰地显示了爆轰波阵面的诱导激波、横波及化学反应区。当爆轰波在左尖点处绕射,受稀疏波作用,诱导激波与化学反应区明显分离,导致爆轰波衰减为爆燃。分离的诱导激波和折皱的化学反应区在纹影图上清晰可见。诱导激波在垂直支管右壁面反射,诱导二次起爆。畸变爆轰波在水平和垂直支管中均发生马赫反射。提高初压,爆轰波受分叉口几何属性的影响减小,畸变爆轰波在水平和垂直支管下游较易恢复为自持爆轰波。     

Investigation on onset of shock-induced separation

A great number of experimental data indicating shock wave/boundary layer interactions in internal or external supersonic flows were reviewed to make clear the mechanism of the interaction and to decide the onset of shock-induced separation. The interesting conclusions were obtained for the considerably wide range of flow geometries that the onset of separation is independent of the flow geometries and the boundary layer Reynolds number. It is found that the pressure rise necessary to separate the boundary layer in supersonic external flows could be applied to such internal flows as overexpanded nozzles or diffusers. This is due to the fact that the separation phenomenon caused by shock wave/boundary layer interactions is processed through a supersonic deceleration. The shock-induced separation in almost all of interacting flow fields is governed by the concept of free interaction, and the onset of shock-induced separation is only a function of the Mach number just upstream of shock wave. However, physical scales of the produced separation are not independent of the downstream flow fields.     

激波与转捩边界层干扰非定常特性数值分析

激波与边界层干扰的非定常问题是高速飞行器气动设计中基础研究内容之一.以往研究主要针对层流和湍流干扰,在分离激波低频振荡及其内在机理方面存在着上游机制和下游机制两类截然不同的理论解释.分析激波与转捩边界层干扰下非定常运动现象有助于进一步加深理解边界层状态以及分离泡结构对低频振荡特性的影响规律,为揭示其产生机理指出新的方向.采用直接数值模拟方法对来流马赫数2.9,24?压缩拐角内激波与转捩边界层干扰下激波的非定常运动特性进行了数值分析.通过在拐角上游平板特定的流向位置添加吹吸扰动激发流动转捩,使得进入拐角的边界层处于转捩初期阶段.在验证了计算程序可靠性的基础上,详细分析了转捩干扰下激波运动的间歇性和振荡特征,着重研究了分离泡展向三维结构对激波振荡特性的影响规律,最后还初步探索了转捩干扰下激波低频振荡产生的物理机制.研究结果表明:分离激波的非定常运动仍存在强间歇性和低频振荡特征,其时间尺度约为上游无干扰区内脉动信号特征尺度的10倍量级;分离泡展向三维结构不会对分离激波的低频振荡特征产生实质影响.依据瞬态脉动流场的低通滤波结果,转捩干扰下激波低频振荡的诱因来源于拐角干扰区下游,与流场中分离泡的收缩/膨胀运动存在一定的关联.     

Wave induced thermal boundary layers in a compressible fluid: analysis and numerical simulations

The response of a semi-infinite compressible fluid to a step-wise change in temperature of its boundary is investigated analytically and numerically. Numerical results of the boundary layer structure are compared with Clarke’s analytical solution for a gas with thermal conductivity proportional to temperature. To avoid unwanted numerical dissipation in the numerical analysis, the space-time conservation element and solution element (CESE) method has been adopted to solve the unsteady 1-D Navier-Stokes equations. Good agreement between analytical and numerical results has been found for the development of the thermal boundary layer on a long time scale. Weak shock waves and expansion waves induced by the thermal boundary layer due to its compressibility, are observed in the numerical simulation. Finally, the numerical method has been applied to the reflection of a non-linear expansion wave and to a shock wave from an isothermal wall, thereby illustrating the effect of the boundary layer on the external flow field.     

Computational analysis of impinging shock-wave boundary layer interaction under conditions of incipient separation

The interaction of an oblique shock wave with a turbulent boundary layer under conditions of incipient separation is analyzed by means of large-eddy simulation (LES) and Reynolds-averaged Navier–Stokes (RANS) turbulence models, with the objective to explore their predictive capabilities, in particular with respect to the unsteady features of the interaction. Consistent with earlier direct numerical simulations, we have found that the flow dynamics in the interaction zone is characterized by strong intermittency associated with the formation of scattered spots of flow reversal near the nominal position of the reflected shock. Comparison with experimental results (at much larger Reynolds number) show that the qualitative features of the interaction are predicted reasonably well by both LES and RANS models. RANS models supplemented with a semi-empirical closure are also found to provide reasonable estimate of the fluctuating pressure loads at the wall.     

Incipient separation in shock wave/boundary layer interactions as induced by sharp fin

The incipient separation induced by the shock wave/turbulent boundary layer interaction at the sharp fin is the subject of the present study. Existing theories for the prediction of incipient separation, such as those put forward by McCabe (1966) and Dou and Deng (1992), can thus far only predic the direction of surface streamline and tend to overpredict the incipient separation condition based on the Stanbrook’s criterion. In this paper, the incipient separation is first predicted with Dou and Deng (1992)’s theory and then compared with Lu and Settles’ (1990) experimental data. The physical mechanism of the incipient separation as induced by the shock wave/turbulent boundary layer interactions at sharp fin is explained via surface flow pattern analysis. Furthermore, the reason for the observed discrepancy between the predicted and experimental incipient separation conditions is clarified. It is found that when the wall-limiting streamlines behind the shock wave becomes aligned with one ray from the virtual origin as the strength of the shock wave increases, the incipient separation line is formed at which the wall-limiting streamline becomes perpendicular to the local pressure gradient. The formation of this incipient separation line is the beginning of the separation process. The effects of Reynolds number and Mach number on incipient separation are also discussed. Finally, a correlation for the correction of the incipient separation angle as predicted by the theory is also given.

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Three dimensional shock wave/boundary layer interactions

An investigation into a three-dimensional, curved shock wave interacting with a three-dimensional, curved boundary layer on a slender body is presented. Three different nose profiles mounted on a cylindrical body were tested in a supersonic wind tunnel and numerically simulated by solving the Navier–Stokes equations. The conical and hemispherical nose profiles tested were found to generate shock waves of sufficient strength to separate the boundary layer on the cylinder, while the shock wave generated by the ogival profile did not separate the boundary layer. For the separated flow, separation was found to occur predominantly on the windward side of the cylinder with the lee-side remaining shielded from the direct impact of the incident shock wave. A thickening of the boundary layer on the lee-side of all the profiles was observed, and in the conical and hemispherical cases this leads to the re-formation of the incident shock wave some distance away from the surface of the cylinder. A complex reflection pattern off the shock wave/boundary layer interaction (SWBLI) was also identified for the separated flow cases. For comparative purposes, an inviscid simulation was performed using the hemispherical profile. Significant differences between the viscous and inviscid results were noted including the absence of a boundary layer leading to a simplified shock wave reflection pattern forming. The behaviour of the incident shock wave on the lee-side of the cylinder was also affected with the shock wave amalgamating on the surface of the cylinder instead of away from the surface as per the viscous case. Test data from the wind tunnel identified two separation lines present on the cylindrical surface of the hemispherical SWBLI generator. The pair of lines were not explicitly evident in the original CFD simulations run, but were later identified in a high-resolution simulation.     

Numerical computations of supersonic inlet flow

Air‐breathing propulsion systems for high‐speed space travel applications are studied. Ramjets and scramjets have been identified as potential candidates. The flow inlets of such systems are modelled with a simulation that can predict all complex inlet flow features, including shock due to forebody, multiple shock reflections, normal shock, shock–boundary layer interaction and associated separation for two‐dimensional and axisymmetric inlets. Computed values are in good agreement with experimental data. Copyright © 2001 John Wiley & Sons, Ltd.     

数值研究平板方舵激波-湍流边界层干扰

数值研究了平板方舵激波-湍流边界层干扰流场。模拟出了分离激波与弓型激波砬撞后形成的“λ”激波结构;消晰地显示了分离区中的旋涡结构,发现流场中会出现二次分离涡,并从理论上分析了流场对称面涡心形态与非定常的关系,得到了涡心为不稳定螺旋点或出现极限环是非定常流动特征的新结论。     

HLLE++格式在高马赫数空腔流动模拟中的应用

空腔流动存在剪切层运动、涡脱落与破裂,以及激波与激波、激波与剪切层、激波与膨胀波和激波/涡/剪切层相互干扰等现象,流动非常复杂,特别是高马赫数(M>2)时,剪切层和激波更强,激波与激波干扰更严重,对数值格式的要求更高,既需要格式耗散小,对分离涡等有很高的模拟精度,又需要格式在激波附近具有较大的耗散,可以很好地捕捉激波,防止非物理解的出现。Roe和HLLC等近似Riemann解格式在高马赫数强激波处可能会出现红玉现象,而HLLE++格式大大改善了这种缺陷,在捕捉高超声速激波时避免了红玉现象的发生,同时还保持在光滑区域的低数值耗散特性。本文在结构网格下HLLE++格式的基础上,通过改进激波探测的求解,建立了基于非结构混合网格的HLLE++计算方法,通过无粘斜坡算例,验证了HLLE++格式模拟高马赫数流动的能力,并应用于高马赫数空腔流动的数值模拟,开展了网格和湍流模型影响研究,验证了方法模拟高马赫数空腔流动的可靠性和有效性。     

Progress of the unified wave-particle methods for non-equilibrium flows from continuum to rarefied regimes

We utilize the nonlinear acoustic solver (NLAS) and Ffowcs-Williams/Hawkings (FW-H) equation to investigate the noise generation and radiation due to shock (wave) and boundary layer interaction (SBLI) in the inlet duct. A classical benchmark for SBLI is chosen to validate the flow features and numerical results show good agreement with experimental results. In the simulation of the noise generated by SBLI, the inlet buzz phenomenon is successfully observed. The oscillation of the normal shock is a kind of little buzz and the oscillation of inner shocks is a kind of big buzz with a frequency around 100 Hz. In the far-field, frequency spectrums show a dominant frequency close to the frequency of inner shocks oscillation. This indicates that the oscillation of inner shocks determines the magnitude of the overall sound pressure level (OASPL) of the far-field noise.