高精度格式WCNS-E-5计算物面热流

采用加权紧致非线性格式WCNS-E-5,四阶精度的二阶导数差分近似以及四阶精度的边界格式构造了高精度算法,对高超声速粘性流动的物面热流进行数值研究.首先考察了壁面网格雷诺数对驻点热流的影响,然后开展了边界格式对热流计算结果影响的研究,最后对大攻角钝锥绕流进行了数值模拟.研究结果表明:WCNS-E-5能降低边界层内网格分辨率,全场高精度的WCNS-E-5计算得到的流场图像清晰、真实、分辨率高,热流值准确、可靠.     

高精度非定常激波装配法

为正确模拟高超声速绕流中,来流小扰动与弓形激波之间的干扰对流动特征的影响,将弓形激波作为动边界,利用非定常特征关系处理激波处的边界条件.应用五阶精度迎风紧致格式和六阶精度的对称格式与三阶精度的R-K方法相结合,建立高精度非定常激波装配方法.采用该方法数值模拟钝锥高超声速定常流场和二维抛物外形高超声速边界层流动的感受性问题,数值模拟来流小扰动与弓形激波干扰激波后非定常扰动流场,研究扰动波进入边界层产生边界层不稳定波的特征.     

湍流模式对转捩的初步研究

采用二阶时间精度的LU-SGS-τTS目的进行时间推进求解Navier-Stokes方程组,用基于Roe的六阶对称TVD格式离散对流项,并采用Baldwin-Lomax和Spalart-Allmaras模式对平板边界层、Aerospatial A-profile翼型低速大迎角及RAE-2822翼型跨声速激波/边界层干扰流动进行数值分析,自动捕捉流动的转捩过程,获取了有意义的结果.     

一类新型激波捕捉格式的耗散性与稳定性分析

高超声速流动是高复杂性的可压缩黏性流动, 其中存在激波、剪切层、激波/激波干扰、激波/边界层干扰、旋涡与分离流动等复杂流场结构. 对其进行准确模拟需要使用低耗散、强鲁棒性的激波捕捉方法. 本文基于一类新型的通量项分裂方法, 提出了一种耗散低且鲁棒性好的激波捕捉格式K-CUSP-X. 对该格式的耗散性和激波稳定性进行了详细的理论分析, 得到了格式激波稳定的数值条件. 推论认为, 迎风格式激波稳定的充分条件为速度扰动量具有衰减性, 数值实验验证了该推论. 研究表明, 该格式与Toro提出的通量分裂格式K-CUSP-T相比, 在保证精确捕捉接触间断的同时, 又具有更好的稳定性, 在激波处不会产生“红玉”现象.     

再入体头部粘性流场高分辨率格式数值模拟

TVD格式是目前数值研究以激波为主要特征之一的超声速、高超声速流场的最先进的算法之一。本文用二阶迎风TVD格式,对三种烧蚀外形的轴对称粘性流场和10°钝锥有攻角三维粘性流场进行了数值模拟,得到了高质量的头部脱体激波和与实验结果及直线推进法计算一致的物面压力分布,表明了TVD格式在再入体粘性绕流计算中的独特优势。     

高超声速层流绕三维压缩拐角的数值模拟

通过数值模拟研究了高超声速来流绕过压缩拐角的层流分离三维流动特性.数值方法采用三维N-S方程,结合2阶精度Roe格式以及分区结构网格有限体积法进行离散.数值模拟的空间激波结构与实验纹影结果符合较好;激波/边界层干扰区内3条纵向线上的计算压力分布与实验结果进行了对比分析,计算获得在三维楔侧面存在低压力区,与实验结果反映的规律一致,计算结果表明低压力区是由楔体侧缘尖端发起的二次涡的抽吸作用造成的.此外,在楔体后端尾流区的低压沿边界层内的亚声速区往上游传递了一定距离.      

Schardin问题的数值研究

激波绕过三角楔(Schardin问题)时会产生激波马赫反射与绕射、 三角楔尾涡与涡串等复杂物理现象. 本文利用三阶精度加权基本无振荡(WENO)格式、 结构化矩形网格的自适应加密方法与沉浸边界法对Schardin问题进行了数值模拟. 数值结果清晰地显示了激波与三角楔相互作用, 在楔面发生马赫反射以及在楔角绕射诱导主涡的过程, 并与Schardin等的实验结果及相关数值结果完全符合. 另外, 数值结果还详细反映了先前实验与数值结果没有详细讨论的主涡滑移层上的涡串生成机理, 以及激波与涡串相互作用和产生声波的     

Schardin问题的数值研究

激波绕过三角楔(Schardin问题)时会产生激波马赫反射与绕射、 三角楔尾涡与涡串等复杂物理现象. 本文利用三阶精度加权基本无振荡(WENO)格式、 结构化矩形网格的自适应加密方法与沉浸边界法对Schardin问题进行了数值模拟. 数值结果清晰地显示了激波与三角楔相互作用, 在楔面发生马赫反射以及在楔角绕射诱导主涡的过程, 并与Schardin等的实验结果及相关数值结果完全符合. 另外, 数值结果还详细反映了先前实验与数值结果没有详细讨论的主涡滑移层上的涡串生成机理, 以及激波与涡串相互作用和产生声波的过程.     

激波与SF6球形气泡相互作用的数值研究

本文基于大涡模拟方法, 采用高阶精度格式对平面入射激波以及不同反射距离条件下的反射激波与SF₆重气泡相互作用过程进行了三维数值模拟. 数值结果清晰地显示了SF₆重气泡在激波作用下诱导Richtmyer-Meshkov不稳定性过程, 揭示了入射激波以及反射激波在气泡界面聚焦诱导射流的过程, 详细分析了不同反射距离条件下反射激波与SF₆重气泡作用过程及流场结构.     

三角翼大迎角非对称旋涡破裂数值模拟

采用由伪时间子迭代格式实现的二阶精度LU SGS方法进行时间推进,并以Jameson中心加人工粘性格式进行空间离散,应用层流假设和Baldwin Lomax(B-L)模式,求解雷诺平均的薄层Navier Stokes(N-S)方程组以模拟细长三角翼大迎角流动.为验证本方法,首先计算了具备实验数据的65°后掠角三角翼大迎角流场,获得令人满意的结果;然后模拟了80°后掠角的细长三角翼,研究其大迎角非对称旋涡破裂特性.     

Controlled Vortex Breakdown on Modified Delta Wings

This paper studies the effect of perturbation to the breakdown of the leading-edge vortices over delta wings. The passive perturbation in the normal direction is achieved by installing the hemisphere-like bulges on the delta wing along the projection of the vortices. The key purpose of this perturbation is to delay or suppress vortex breakdown over delta wings according to the self-induction mechanism theory. The design of bulge-like surface for delta wings offers a minimization of initial vorticity gradient and an elimination of linearly mutual induction within the vortex core. Three delta wings with swept angles of 60°, 65dg and 70° have been used. Dye flow visualization and force measurement in different water tunnels are performed at the water speed of U=0.10, 0.15, 0.20 and 0.25 m/s. In flow visualization, the results show contributions of bulges as perturbation to leading-edge vortices. The best outcome of perturbing the vortex core occurs in the case of the 65° delta wing. The breakdown positions on the 65° delta wing are delayed in almost the entire range of angles of attack, and that, the results are presented here.     

Quantitative visualization of the leading-edge vortices on a delta wing by using pressure-sensitive paint

Leading-edge vortices on a simple delta wing were visualized by using pressure-sensitive paint (PSP). PSP is an optical pressure measurement technique based on oxygen quenching of luminescent molecules. In the present study, we used PSP composed of platinum octaethylporphyrine (PtOEP) and fluoropolymer (poly-IBM-co-TFEM [Poly (isobutylmethacrtlate-co-trifluoroethylate)]). This new paint has higher sensitivity to pressure and lower sensitivity to temperature than previous ones, reducing an error due to temperature variation during a wind tunnel test. A thin coating of PSP was applied to a delta wing model with 70-degree leading-edge sweep. The coating was excited by Xenon light and emission from the coating was detected by a high-resolution CCD camera. Tests were done at subsonic speeds in the 0.2-m Supersonic Wind Tunnel at the National Aerospace Laboratory in Japan. Complicated flow structures on the delta wing including primary and secondary vortices were clearly visualized using pressure-sensitive paint. An a priori calibration technique was used to convert measured luminescent intensity into pressure. The obtained pressure distributions were in good agreement with pressure tap data. Pressure maps were obtained for various Mach numbers, Reynolds numbers and angles of attack. It was found that an increase in Mach number delayed vortex breakdown while Reynolds number had little effect on the vortex formation.     

The effects of vortex breakdown on the static rolling aerodynamics of finned slender body

The static rolling aerodynamics of a finned slender body is numerically studied in this paper.Simulation results show a nonlinear uprising of the rolling moment when the angle of attack is greater than 20°in subsonic flows.Asymmetric vortex break down phenomenon on the"horizontal"rudders is found to be responsible for this phenomenon.By introducing the geometric-equivalent angle of attack and geometric-equivalent sweep angle,the cause of the nonlinear rolling moment characteristics can be explained by the delta wing vortex breakdown analysis.     

Flow field measurements of leading-edge separation vortex formed on a delta wing with vortex flaps

Wind tunnel tests are carried out using a 70 delta wing model with leading-edge vortex flaps. The structure of the leading-edge separation vortex over the leading-edge vortex flap is measured by use of a 5 holes pitot probe, surface pressure measurement technique and oil flow visualization technique. Separation vortices formed on a plain delta wing, on a vortex flap and inboard the vortex flap hinge line are clearly visualized. Results indicate that the flow around the vortex flaps is classified into several different cross flow patterns. The streamwise flap deflection angle is defined to discuss the vortex flap performance. The optimum lift to drag ratio is attained when the amount of the wing angle of attack is not far different from that of the streamwise flap deflection angle, as long as the vortex flap is deflected modestly.     

基于非线性涡格法的快速静气动弹性数值模拟技术

基于2.5D RANS数据和VLM耦合的方式,发展了一种考虑非线性流动效应的混合型涡格法HVLM.采用矩形直/后掠机翼两个外形的跨声速算例,通过与VLM、三维CFD计算数据的比较,对HVLM的气动力预测精度进行了分析与评估.对比结果表明,HVLM在大幅降低时间成本的前提下可以获得和三维CFD方法预测值十分接近的计算数据,对线化VLM方法的修正效果显著.然后,HVLM与悬臂梁有限元求解耦合,实现了一种面向三维机翼的快速静气动弹性数值模拟技术,并通过矩形直机翼算例进行了验证.耦合算例的时间分析数据表明,HVLM/Beam耦合的方式能够在10 s以内完成1次三维机翼静气动弹性分析,在气动/结构耦合分析、优化设计方面展示出了良好的应用前景.      

The impact of the pitching motion on the structure of the vortical flow over a slender delta wing under sideslip angle

The primary purpose of this investigation is to observe the effect of the pitching motion on the vortical flow structure and bursting of leading-edge vortices over a delta wing under the sideslip angle, β using a dye visualization technique. In the current work, a delta wing with a sweep angle of Λ = 70° was oscillated in upstroke and downstroke direction to be able to discover the influence of pitching motion on the flow characteristics of the delta wing. The values of mean angles of attack were selected as α_m = 25° and α_m = 35°, and the sideslip angle was altered from β = 0 to 16°. The delta wing oscillated with the various periods of T_e = 5 s, 20 s, and 60 s, respectively. Amplitude of motion was adjusted as α_o =  ± 5°. It is found that the pitching motion of the delta wing under the sideslip angle β varies the location of the vortex bursting and vortical flow structure substantially.

     

双三角翼非定常俯仰运动实验与数值模拟

文章针对双三角翼大振幅正弦俯仰运动过程中的非定常载荷和流动特性开展了实验与数值模拟研究,并与相同主翼后掠角的单三角翼进行了对比.实验研究在低速回流式水槽中开展,所采用的实验模型为边条后掠角为75°,主翼后掠角为50°的双三角翼全模,俯仰运动的旋转轴位于主翼弦长的2/3处,振幅为0~60°,运动的缩减频率k=0.03,0.06,0.12,0.24,0.48.实验Reynolds数以主翼弦长为参考Re=1.69×104.在水槽的测力实验中,发现非定常流动力的迟滞现象,并且随着非定常运动缩减频率的增大,流动的迟滞效应也随之增大.与相同主翼后掠角的单三角翼相比,双三角翼的迟滞环在低缩减频率下更小,但随着缩减频率的增大,这种差距逐渐减小.在数值模拟研究中,采用DDES湍流模型对俯仰双三角翼的流场进行了数值模拟.流场结果表明,在较低的缩减频率下,主翼吸力面的前缘涡是影响气动力的主要因素,非定常流动力的迟滞效应主要与前缘涡在上仰过程中的延迟破裂和下俯过程中的延迟恢复有关;在较高的缩减频率下,机翼前缘涡对气动力的影响减小,由机翼俯仰角速度而产生的环量力成为了气动力的主导因素,因此在较高缩减频率下,单三角翼与双三角翼的升力特性趋于一致.      

基于雷诺应力模型的高精度分离涡模拟方法

基于Speziale-Sarkar-Gatski/Launder-Reece-Rodi(SSG/LRR)-ω雷诺应力模型发展了一类分离涡模拟方法,结合高精度加权紧致非线性格式在典型翼型及三角翼算例中进行了验证,并和传统基于线性涡粘模型的分离涡模拟方法进行了对比.结果表明:基于SSG/LRR-ω模型的分离涡模拟方法,提高了原雷诺应力模型对非定常分离湍流的模拟能力;同时相比于传统基于线性涡粘模型的分离涡模拟方法,尤其是在翼型最大升力迎角和三角翼涡破裂迎角附近,该方法在平均气动力预测的准确度、分离湍流模拟的精细度等方面更加优秀.