超声速边界层内小激波的数值模拟

本文通过有限差分计算求解了NS方程,对超声速二维平板边界层的非线性稳定性问题进行了模拟.流场的来流马赫数为4.5.平板从尖前缘开始,流场在前缘处产生一道前缘激波.数值计算模拟了边界层对成倍频关系的双频扰动波的吸收问题.通过计算发现,在一定频率下边界层内可出现小激波.计算给出了小激波的发展过程.     

高超声速边界层感受性问题高精度数值模拟

通过五阶WENO格式和六阶对称紧致格式以及三阶TVD R-K结合的方法,对存在强激波和小扰动相互干扰的高超声速边界层感受性问题进行研究.结果表明:此方法能够模拟边界层内不稳定波的产生和发展以及小扰动和激波相互干扰等现象,因此能广泛应用于含激波的感受性等问题的可压缩湍流直接数值模拟以及具有激波和边界层干扰等复杂流场的计算.     

高背压进气道中内外流耦合作用的大涡模拟

采用大涡模拟研究了出口堵塞比为50.8%的轴对称进气道流动,重点考察了内外流耦合作用下流动的非定常特性。采用国家数值风洞（NNW）工程仿真软件进行数值模拟,得到的壁面平均压力、瞬时压力分布与试验数据符合良好。分析表明：为匹配出口背压,进气道在喉道区域形成激波串结构,使内流道流场分为上游超声速区、中部激波串区以及下游亚声速区;在激波串区,剧烈的逆压梯度产生了分离激波、激波串、分离区及分离剪切层等复杂结构;伴随着激波串运动和边界层大尺度分离,进气道壁面压力出现宽频脉动特征。脉动压力的时空分布表明：内流道脉动压力以扰动波的形式传播,为此建立的声反馈模型能较好地预测亚声速区的主导频率。相关性分析表明：激波串运动受上下游流动耦合作用,其中,频率为St=0.7的运动主要受上游流动影响,频率为St=0.9的运动主要受下游压力扰动波影响。     

MVG控制斜激波/湍流边界层干涉的大涡模拟

采用浸没边界法(IBM)对带有微型涡发生器(MVG)控制器的激波/湍流边界层干涉流动进行了大涡模拟(LES)。以来流马赫数为2.3的斜激波(由平板上方8°楔产生)为基本流动入射平板湍流边界层,通过在干涉区前布置MVG阵列来控制激波诱导的边界层分离。采用浸没边界法处理MVG的复杂几何,分析了MVG尾迹区平均流速度剖面,雷诺应力,瞬态旋涡结构。结果表明:时均流场显示MVG尾迹区存在一对对转的主流向涡,流向涡加剧了边界内的动量交换从而增加了边界层抗分离能力,而瞬态流场则反映出MVG尾迹区的剪切层由于Kelvin-Helmholtz(K-H)不稳定性会卷起为一列展向旋涡。     

超声速膨胀角入射激波/湍流边界层干扰直接数值模拟

为了揭示膨胀效应对激波/湍流边界层干扰区内复杂流动现象的影响规律,采用直接数值模拟方法对来流马赫数2.9、30°激波角的入射激波与10°膨胀角湍流边界层相互作用问题进行了数值研究。系统地探讨了激波入射点分别位于膨胀角上游、膨胀角角点和膨胀角下游3种工况下膨胀角干扰区内若干基本流动现象,如分离泡、物面压力脉动及激波非定常运动、湍流边界层统计特性和相干结构动力学过程等。结果表明,激波入射点流向位置改变对分离区流向和法向尺度的影响显著,尤其是当激波入射点位于角点及其下游区域。研究发现,膨胀角干扰区内物面压力脉动强度急剧减小,分离区内压力波向下游传播速度将降低而在膨胀区内将升高,膨胀效应极大地抑制了分离激波的低频振荡运动。相较于入射激波与平板湍流边界层干扰,入射激波流向位置改变对膨胀角再附区速度剖面对数区及尾迹区影响显著,将导致其内层结构参数升高而外层降低,近壁区内将呈现远离一组元湍流状态的趋势。此外,流向速度脉动场本征正交分解分析指出,主模态空间结构集中在分离激波及剪切层根部附近而高阶模态以边界层内小尺度正负交替脉动结构为主。低阶重构流场结果表明,前者对应为分离泡低频膨胀/收缩过程而后者表征为分离泡高频脉动。     

局部扰动对平板边界层流动稳定性影响的研究

在壁面上构建局部扰动,数值研究不同特征的局部扰动对平板边界层流动稳定性影响的作用机理,进一步探讨壁面扰动的分布、类型、强度大小对三维扰动波在平板边界层流动中的非线性演化规律以及失稳机制的问题.确定什么样的局部扰动能使三维扰动波快速增长或者是什么样的局部扰动能对流动起抑制的作用,同时寻求在三维扰动波向下游发展过程中的传播方向会发生怎样的变化等.对该问题的深入研究,将使人们能深入理解认识平板边界层流动过程中的转捩发生、流动失稳及湍流形成的理论机制.     

高超声速流场对自由流扰动波响应及边界层扰动波演化

采用高阶精度有限差分方法直接数值模拟了来流慢声波作用下的钝锥高超声速绕流瞬态流场,分析了自由流扰动波与流场的相互干扰,并应用Fourier频谱分析了边界层扰动波时域和空域演化.结果显示:弓形激波出现连续的“∽”变形,初始扰动被显著放大,边界层内外扰动模态存在明显差别.从空域(沿流线)演化来看,在球头附近低频扰动为主导模态,出了球头区,总扰动模态中的低频和高频成分比例迅速转变,高频模态成分显著地增大.至流场下游,大部分低频分量衰减,或者不再增长,边界层内仅存在特殊频率的不稳定波快速增长.从时域演化来看,比起其他模态,主导模态的发展对上游激励的依赖更大.无论时域还是空域演化,都存在模态竞争现象.      

超声速边界层三维扰动引起小激波的数值研究

通过直接数值模拟的方法,研究了M=4.5的超声速边界层中三维扰动的演化。以某一剖面为入口,加入一个及一对三维T-S波,发现随展向波数的增加,扰动幅值的增长率逐渐减小,证实了M=4.5的超声速边界层中,当三维扰动达到一定幅值时会有小激波出现,为建立可压缩流动稳定性理论提供了依据。     

高超声速激波/边界层干扰流动数值模拟研究

在基于Roe格式的全Navier-Stokes方程计算流体力学(CFD)代码中,发展了一种局部熵修正方法,克服了传统熵修正方法在边界层流动模拟中耗散过大的缺点,可用于更加准确的模拟激波/边界层干扰的复杂高超声速流动。对典型高超声速双锥边界层分离与激波干扰的复杂流动进行了模拟,研究了网格密度和熵修正方法耗散性对计算结果的影响。研究表明:高超声速双锥边界层分离与激波干扰流动的数值模拟结果对网格具有很强的敏感性,过稀的网格将产生严重的分离流动预测偏差;低耗散性的局部熵修正方法能更加准确地模拟复杂的高超声速激波与边界层分离流动干扰现象。     

压缩-膨胀湍流边界层平均摩阻分解

采用直接数值模拟对来流马赫数2.9、24°压缩-膨胀折角构型中激波与湍流边界层干扰问题进行了研究。重点关注膨胀折角法向高度对激波干扰区以及下游平板边界层流动的影响。研究发现，当高度足够大时，激波干扰区内未受下游膨胀波的影响，此时的流动特征与传统的压缩折角干扰构型一致。高度较小时，脱体剪切层的再附过程受到下游膨胀波的加速影响，导致再附点向上游移动，分离泡发生剧烈收缩。对上、下游平板湍流边界层应用了平均摩阻分解技术，比较了湍流边界层在平衡和非平衡状态下的差异。分析发现，膨胀折角区域的高摩阻现象主要与摩阻分解后的C_f1项与C_f3项相关。高度变化对C_f1项影响较小，而对C_f2项影响显著。高度变化体现在：下游平板上G9rtler涡结构强度以及层流化现象对C_f2项贡献的差异。     

Self-sustained shock oscillations on airfoils at transonic speeds

Self-sustained shock wave oscillations on airfoils at transonic flow conditions are associated with the phenomenon of buffeting. The physical mechanisms of the periodic shock motion are not yet fully understood even though experiments performed over fifty years ago have demonstrated the presence of oscillatory shock waves on the airfoil surfaces at high subsonic speeds. The unsteady pressure fluctuations generated by the low-frequency large-amplitude shock motions are highly undesirable from the structural integrity and aircraft maneuverability point of view. For modern supercritical wing design with thick profiles, the shock-induced fluctuations are particularly severe and methods to reduce the shock wave amplitudes to lower values or even to delay the oscillations to higher Mach numbers or incidence angles will result in expanding the buffet boundary of the airfoil. This review begins with a recapitulation of the classical work on shock-induced bubble separation and trailing edge separation of a turbulent boundary layer. The characteristics of the unsteady pressure fluctuations are used to classify the types of shock-boundary layer interaction. The various modes of shock wave motion for different flow conditions and airfoil configurations are described. The buffet boundaries obtained using the standard trailing edge pressure divergence technique and an alternative approach of measuring the divergence of normal fluctuating forces are compared to show the equivalence. The mechanisms of self-sustained shock oscillations are discussed for symmetrical circular-arc airfoils at zero incidence and for supercritical airfoils at high incidence angles with fully separated flows. The properties of disturbances in the wake are examined from linear stability analysis of two-dimensional compressible flows. The advances in high-speed computing make predictions of buffeting flows possible. Navier–Stokes solvers and approximate boundary layer-inviscid flow interaction methods are shown to give good correlation of frequencies and other unsteady flow characteristics with experiments. Finally, passive and active methods of shock oscillation control show promising results in delaying buffet onset to higher Mach numbers or incidence angles, thus enhancing the transonic performance of airfoils.     

超声速压气机叶栅前缘通道激波损失的鼓包控制研究

为了有效减小超声速压气机叶栅变进气马赫数条件下的前缘通道激波损失及由激波诱导的边界层分离,提出了一种带有平直过渡区的新型鼓包结构,并采用数值方法详细分析了新型鼓包结构对激波与激波/边界层相互作用机理以及鼓包几何尺寸与位置对控制效果的影响机制。研究结果表明:新型鼓包在迎风侧凹面产生的压缩波系有效削弱了前缘通道激波的强度,鼓包过渡区产生的膨胀波系使边界层流体加速,明显抑制了局部流动分离,并使分离提前再附。当某一超声速压气机叶栅的前缘通道激波入射在鼓包的过渡区范围内,鼓包高度为0.35倍的边界层厚度且鼓包迎风侧与背风侧长度分别为过渡区长度4倍与5倍时,可以实现较好的控制效果。此外,与无鼓包方案相比,新型鼓包结构可使超声压气机叶栅在设计工况下的总压损失减少4.6%,同时超声速压气机叶栅进气马赫数在1.65~1.8范围内仍能取得较好的气动减损效果。      

基于LES方法的平板非定常激波/湍流边界层干扰研究

以高超声速发动机进气道湍流分离控制为应用背景,采用大涡模拟(LES)方法进行马赫数为3.0(唇口附近马赫数约为3.0)的激波/湍流边界层干扰(SWTBLI)流场机理研究.利用扰动循环引入的方法,先得到充分发展湍流场,然后根据斜激波关系式引入激波的方法进行激波/湍流干扰模拟.研究结果显示:充分发展湍流场在激波作用下产生逆...     

等离子体气动激励控制超声速边界层分离的实验研究

等离子体气动激励与超声速气流相互作用已成为高速流动控制领域的研究热点。激波与边界层相互作用现象广泛存在于超声速飞行器之中。本文进行了等离子体气动激励控制压缩角区和激波诱导边界层分离的实验,通过流场纹影显示和壁面静压测量,研究等离子体气动激励如何影响激波、激波如何影响边界层特性的科学问题。实验结果表明:施加毫秒量级表面电弧放电能够前移压缩角区的诱导斜激波,使分离区后移,分离区域增加,但激波强度减弱,流场总压增加;施加微秒量级表面电弧放电能够抑制激波诱导边界层分离,使分离区减小,流场总压减小。基于实验结果,认为毫秒量级表面电弧放电激励控制超声速气流的主要机理为放电过程的焦耳热效应;微秒量级表面电弧放电激励控制超声速气流的主要机理为焦耳热效应和冲击波效应共同作用。     

新型桨尖抑制旋翼跨声速特性的影响分析

为研究不同形式的新型桨尖在抑制旋翼跨声速特性方面的作用,开展了多种桨尖对旋翼局部流动及气动特性影响的数值分析研究.发展了基于高效嵌套网格方法的旋翼流场高精度CFD求解方法.在此基础上,详细分析了桨尖外形对旋翼桨叶跨声速区域激波强度、激波诱导气流分离、桨尖涡尾迹及气动性能的影响.数值结果表明:桨尖的后掠和上反在缓解旋翼跨声速特性方面的作用相对较小;桨尖前掠和下反能更有效地减少桨尖外端跨声速区域,降低该位置激波强度并缓解激波-附面层干扰诱导的气流分离;后掠桨尖在减小旋翼反扭矩方面的整体效果良好,直线前掠桨尖在大桨盘拉力状态能够更有效降低旋翼扭矩(直线前掠30°时,扭矩降低达12.3%),桨尖下反可以有效抑制桨尖涡强度(抛物下反30°时,桨尖涡强度降低50%),并加快桨尖涡尾迹的耗散.      

高精度数值方法在DNS中的应用

利用不同格式模拟扰动波在平板边界层中的传播,考察了格式对T-S波传播过程计算结果的影响。研究了不同计算格式对自由剪切层扰动波增长的影响,结果表明:对于二维扰动波模型,采用二阶精度格式,扰动波在平板边界层中只能出现衰减的结果;在自由剪切层中,虽然能模拟出扰动波在传播过程中幅值的增长,但却在相当长的计算范围内看不见三维波的出现。但对于同样的扰动模型,如果采用高精度格式,在边界层中能出现增长的T-S波,并与线性理论非常吻合;在自由剪切层中,扰动波不仅迅速增长,并能在不很长的距离内,明显的看到三维波的出现,并激发出了三维结构。      

Experimental investigation of expansion effect on shock wave boundary layer interaction near a compression ramp

The experiment is conducted to investigate the effect of expansion on the shock wave boundary layer interaction near a compression ramp. The small-angle expansion with an angle degree of 5° occurs at different positions in front of the compression ramp. The particle image velocimetry and flow visualization technology show the flow structures, velocity field, and velocity fluctuation near the compression ramp. The mean pressure distribution, pressure fluctuation, and power spectral density are measured by high-frequency response pressure transducers. The experimental results indicate that the expansion before the compression ramp position affects the shock wave boundary layer interaction to induce a large-scale separation. But the velocity fluctuation and pressure fluctuation are attenuated near the large-scale flow separation region. When the expansion occurs closer to the compression ramp, the expansion has a more significant impact on the flow. The fluctuation of velocity and pressure is significantly attenuated, and the wall pressure rise of the separation point is reduced obviously. And the characteristic low-frequency spectrum signal related to the unsteadiness of the shock wave boundary layer interaction is significantly suppressed. In addition, variation of the separation region scale at different compression angle degrees is distinctive with the effect of expansion.     

高压涡轮转子间隙泄漏流动的非定常特征研究

利用数值方法求解三维非定常雷诺平均Navier-Stokes方程模拟某跨声速高压涡轮流场,研究了某跨声速高压涡轮流场的非定常特征,通过详细分析动静干涉对间隙泄漏流动的影响,进一步明确了泄漏流周期性变化的规律和成因.研究发现:静子尾缘燕尾波的外侧分支外尾波是间隙内部流动结构变化的主要原因,间隙泄漏涡的周期性变化则受外尾波和尾迹的共同影响.外尾波深入转子通道内部周期性经过间隙,在间隙前缘附近产生很强的逆压梯度,使间隙前部流动方向明显改变而产生大范围分离.外尾波导致间隙泄漏流量明显增加并周期性震荡.在静子尾迹和外尾波共同作用下,泄漏涡强度出现波动且涡的位置前后移动,使泄漏涡呈现明显的非定常性.      

Influence of shock wave impinging region on supersonic film cooling

Shock waves can significantly affect the film cooling for supersonic flow and shock waves may have different influence when impinging in different regions. The present study numerically compared the results of shock wave impinging in three different regions and analyzed the effect of impinging region. The shock wave generators were located at x/s = 5, 25, 45 with 4°, 7° and 10° shock wave incidence. The mainstream Mach number was 3.2 and the coolant Mach number was 1.2 or 1.5. The numerical results illustrated that the shock wave impinged in the further upstream region led to a larger high-pressure region and a larger vortex in the boundary layer. Moreover, placing the shock wave generator upstream resulted in the lower mass fraction of coolant in the downstream region. The velocity in the upstream part of the cooling layer was lower than the midstream and downstream part, which resulted in the less ability to resist the shock wave impingement. Therefore, the upstream impingement deteriorated the cooling performance to a greater extent. The study also manifested that the stronger shock wave had a larger effect on supersonic film cooling. Increasing the coolant inlet Mach number can increase the blowing ratio and reduce the mixing, which was of benefit to improve cooling effect.     

Flow characteristics around airfoils near transonic buffet onset conditions

In transonic flow, buffet is a phenomenon of flow instability caused by shock wave/boundary layer interaction and flow separation. The phenomenon is common in transonic flow, and it has serious impact on the structural strength and fatigue life of aircraft. In this paper, three typical airfoils: the supercritical OAT15A, the high-speed symmetrical NACA64A010, and the thin, transonic/supersonic NACA64A204 are selected as the research objects. The flow fields of these airfoils under pre-buffet and buffet onset conditions are simulated by Unsteady Reynolds Averaged Navier-Stokes (URANS) method, and the mode analysis of numerical results is carried out by Dynamic Mode Decomposition (DMD). Qualitative and quantitative analysis of the shock wave motion, shock wave intensity, shock foot bubble and trailing edge separation, and pressure coefficient fluctuation were performed to attain deep insight of transonic buffet flow features of different airfoils near buffet onset conditions. The results of DMD analysis show that the energy proportion of the steady mode of these airfoils decreases dramatically when approaching the buffet onset angle of attack, while the growth rate of the primary mode increases inversely. It was found that at the onset of buffet, there exist different degrees of merging behavior between shock foot bubble and trailing edge separation during one buffet cycle, and the instability of shock wave and separation induced shear layer are closely related to the merging behavior.