变截面激波管中激波运动规律的初步数值研究

利用有限差分方法求解变截面激波管中存在激波的二维流体流动问题。着重研究初始流场为静止和超音速两种情形。分析讨论所得计算结果,并和已有实验结果进行了比较。     

高空模拟试车台扩压器数值分析

采用数值方法对固体火箭发动机主动引射高空模拟试车系统中的扩压器流场进行模拟求解,比较了含与不含Al2O3粒子对扩压器流场的影响.数值计算结果与试验结果符合性较好,表明数值模拟方法可作为扩压器流场计算的一种研究手段.     

三种湍流对Scramjet燃烧室冷态流场的数值模拟

为验证从超音速混合层中发展而来的修正可压缩性(包括膨胀可压缩性和结构可压缩性两部分)的k-ε湍流模型对复杂超音速流场的预测能力,对复杂的超音速燃烧冲压发动机(Scramjet)燃烧室冷态流场进行数值模拟.将修正k-ε模型、标准k-ε模型以及标准k-ω SST模型的计算结果与实验结果进行了对比.结果表明,修正k-ε模型能够较好地预测该复杂流场,对湍流动能的预测改进较大,优于k-ω SST的结果,与实验吻合得较好;壁面压力分布以及速度分布在激波附近修正明显,与实验结果更加吻合.     

有限差分(激波扑获)法对超音速流场计算的应用

激波扑获差分法是计算带激波的超音速流场的一种有效方法。本文采用此法举例计算了超音速气流绕尖劈、菱形翼剖面、园弧翼剖面的流动,通过具体算例对所用差分格式和几种处置边界条件的方法进行分析讨论。所得结果可作为进一步计算复杂三元问题的参考。     

进口有旋及畸变条件下轴流式扩压器的数值分析

本文应用改进的二维椭圆型通用程序对锥形扩压器流场作了计算。计算扩张角2(?)的范围为8°,12°,15.8°和16°,面积比AR 为1.5～9.0左右。计算中应用湍流的有效粘性模型和二方程k—ε模型。扩压器进口气流条件为无旋的,有旋的和轴向畸变的三种。计算结果与试验性能数据的比较是令人满意的。文中还对JT3D涡扇发动机的环形扩压器进行全流场的计算。本方法可提供设计复杂形状扩压器或其他流域的流场分析,性能预报和试验方案。     

用数值模拟方法确定跨音速风洞几何参数

用多块搭接网格技术及雷诺平均N-S方程数值模拟跨音速翼型风洞流场,着重研究了不同驻室高度、不同引射缝高度以及实验段上下壁不同开闭比对风洞马赫数分布的影响,并研究了实验段上下壁干扰对翼型测压实验的影响,通过计算得到了最优的驻室高度、引射缝高度以及最优开闭比。计算结果表明,不同的驻室高度和引射缝高度对实验段马赫数分布有很大影响;而实验段上下壁开闭比对风洞实验段的阻塞度有很大的影响,随着开闭比的逐渐增加,翼型上的激波位置也逐渐向前移动,只有在合适的开闭比下,翼型的激波位置和压力峰值才与无干扰结果一致。研究结果对风洞建设和性能改进有一定的参考意义。     

柱对称抛射体超音速绕流Mach-Zehnder干涉图分析的数值解法

本文把作者在文献[1]中对“含激波等离子体光谱的Abel变换”得到的四条物理结论和求解含激波的Abel变换的数值解法直接推广到了柱对称抛射体超音速绕流的Mach-Zehnder干涉图分析上。文中对实验测得的干涉条纹移动δ(x,z)与超音速流场中各点处的密度改变ρ(r,z)-ρ_0之间的关系得到了与文献[1]类似的三条物理结论,并给出了从观测到的干涉条纹移动δ(x,z)计算流场中各点处的密度改变ρ(r,z)-ρ_0所用的全部变换系数公式。本文利用文献[2]中的一张柱对称抛射体超音速绕流的Mach-Zehnder干涉图作了流场密度计算,得到了好的结果。文中还给出了流场中含有多个激波时的简化算法。     

模拟超音速流场的一种改进CSCM算法

对模拟超音速流场的 CSCM 方法做了修改,给出了一种更有效的简化算法.简化后的算法保持了 CSCM 的优点,减小了计算量,且使结果的分辨率提高.运用此算法模拟了激波管和喷管流场,给出了预期的结果.     

非设计工况下叶片扩压器与离心叶轮间非定常流动的二维数值模拟

以一实验用离心压气机内叶片扩压器与叶轮间耦合流动为研究对象，采用动静部件统一的计算方法，对叶片扩压器与离心叶轮间二维流场在非设计工况下的非定常流动进行了数值模拟，并与无叶扩压器与离心叶轮间定常流场进行比较．与测量结果比较表明，本文的计算结果具有一定的可信度，并由此说明了非设计工况下，动静部件干涉对流动的影响．     

带背景激波系的凹腔流动特性研究

凹腔作为超燃冲压发动机的一种火焰稳定器受到广泛关注,凹腔剪切层与背景激波系的相互作用影响凹腔火焰稳定器的性能。为深入分析背景激波系对凹腔流动的影响,设计了长深比为13.3的闭式凹腔,将凹腔模型前缘激波和风洞上壁面干扰激波作为背景激波系,在Ma=2的直连式风洞中开展了背景激波系与凹腔剪切层的相互作用的试验,采用高速纹影系统对瞬态流场进行了捕捉,重点关注背景激波系和凹腔剪切层的动态变化特性。采用纹影序列的本征正交分解来研究流场中的主要相干结构,采用快速傅里叶变换和连续小波变换对流场的频率域特征进行了分析。结果表明：在背景激波与剪切层相互作用下,激波结构产生大尺度振荡,凹腔内流动结构产生小尺度脉动。通过对激波位置的傅里叶变换分析,发现激波振荡的主导频率集中在90～400 Hz的范围内。通过对纹影图像的空间傅里叶变换分析,发现5 kHz以下的流场振荡主要由激波振荡引起,5 kHz以上的流场脉动主要由凹腔内流动结构引起。     

高超声速流动中双尺度湍流模式的应用

论文研究了双尺度湍流模式,并对其在壁面附的近长度尺度进行了修正,选择四个基准流动－超声速和高超声速二维压缩拐角,锥柱裙组合体绕流和斜激波／平板湍流边界层干扰－进行了数据计算,数值计算和实验结果的比较表明修正双尺度湍流模式对流动分离,摩阻和热流的计算具有更好的效果。     

径流式涡轮排气扩压器流动性能模拟研究

本文研究了径流式涡轮排气扩压器特性对涡轮增压器与中速柴油机匹配性能的影响。在研究过程中，假定排气扩压器的气流为一元准稳定不可压缩气流，试验装置采用压缩气源供气。根据相似理论模拟中速柴油机排气的气流状态进行内壁为直线型面（定为Ａ型）及内壁为曲线型面（定为Ｂ型）这两种型面的扩压器模拟试验。计算和试验结果表明：（１）不论Ａ型或Ｂ型排气扩压器都能改善径流式涡轮的性能，从而提高涡轮效率及涡轮增压器总效率；（２）当扩张角较大时，Ｂ型扩压器因有较高的扩压器效率及压力恢复系数，其性能优于Ａ型扩压器，这将更有利于改善涡轮增压器与中速柴油机的匹配性能。     

无叶扩压器中恒定流的损失预测

本文分析了离心式压缩机无叶扩压器内流扬的特征,并建立了其内部流动的数学物理模型.认为沿无叶扩压器流道中分面具有非对称的粘性流动性质.在非径向流动时,其两侧壁上的边界层产生扭曲,并有一个从进口起始段到下游充分发展段的变化过程.文中着重对无叶扩压器的流动损失规律作了一些探讨,采用了内层变量表示的速度壁面律,以便计及边界层内部温度和径向压力梯度对速度分布和壁面摩擦系数的影响,从而提供了一种预测无叶扩压器性能的新方法.经不同进口条件下的实验验证,证明本文提出的性能预测方法是可行的,可以应用于粘性可压缩非对称流,也可以求解径向流和沿无叶扩压器中分面对称的流动.     

离心式压缩机无叶扩压器内三元紊流附面层的计算

提出了一种计算无叶扩压器两壁面上不可压缩三元紊流附面层的方法。分析中假设径向速度剖面为二阶梯形,且附面层内的速度分布采用能明显表示主流流线偏转角影响的Johns ton的三角形模型。将附面层的动量积分方程与主流的关系式联立起来求解。计算出的无叶扩压器的气动特性与作者的试验结果吻合良好。     

Parametric effects on the combustion flow field of a typical strut-based scramjet combustor

The flame-holding mechanism in hypersonic propulsion technology is the most important factor in prolonging the duration time of hypersonic vehicles.The two-dimensional coupled implicit Reynolds-averaged Navier-Stokes equations,the shear-stress transport k-ω turbulence model and the finite-rate/eddy-dissipation reaction models were used to simulate the combustion flow field of a typical strut-based scramjet combustor.We investigated the effects of the hydrogen-air reaction mechanism and fuel injection temperature and pressure on the parametric distributions in the combustor.The numerical results show qualitative agreement with the experimental data.The hydrogen-air reaction mechanism makes only a slight difference in parametric distributions along the walls of the combustor,and the expansion waves and shock waves exist in the combustor simultaneously.Furthermore,the expansion wave is formed ahead of the shock wave.A transition occurs from the shock wave to the normal shock wave when the injection pressure or temperature increases,and the reaction zone becomes broader.When the injection pressure and temperature both increase,the waves are pushed out of the combustor with subsonic flows.When the waves are generated ahead of the strut,the separation zone is formed in double near the walls of the combustor because of the interaction of the shock wave and the boundary layer.The separation zone becomes smaller and disappears with the disappearance of the shock wave.Because of the horizontal fuel injection,the vorticity is generated near the base face of the strut,and this region is the main origin for turbulent combustion.     

机匣附面层对涡轮转子叶顶流动及换热的影响

以GE-E3型燃气涡轮发动机第1级高压涡轮转子为对象,通过改变进口段长度、机匣的壁面条件以及叶顶间隙的高度,调节二次流与泄漏流之间的强弱关系,分析了机匣附面层对叶顶区域气流流动和叶顶壁面换热特性的影响,并研究了叶顶边缘的倒圆处理对叶顶气流流动和壁面换热的影响.结果表明:泄漏流与二次流的相互作用,导致叶顶头部吸力面侧产生了高换热系数区域;减少二次流或增加泄漏流,均可使得叶顶头部吸力面侧的高换热系数区域减小,压力面侧的高换热系数区域增大.     

壁判据检验激波/层流边界层干扰流动的数值计算

针对二维激波/层流边界层干扰流动数值计算结果的可信度问题,根据高智研究员提出的壁判据对数值计算结果进行了分析和可信度评估.对于二维可压流动,在壁面附近,已经证明壁判据与NS方程解析解相吻合,数值计算结果也表明达到了网格无关性,而利用Opencfd算出的数值解却发现:随着网格的逐渐加密,切向动量壁判据和能量壁判据吻和得越来越好,而法向动量壁判据仍然有较小的差距,但并不影响流场的结果,其原因是壁面压力边界条件选取不当.数值计算结果表明,壁判据可以作为含壁面流动的一种CFD计算结果可信度评估的工具,具体的评估标准仍需进一步研究.     

The experimental study of interaction between shock wave and turbulence

The interaction between shock wave and turbulence has been studied in supersonic turbulent mix layer wind tunnel. The interaction between oblique shock wave and turbulent boundary layer and the influence of large vortex in mix layer on oblique shock wave have been observed by NPLS technique. From NPLS image, not only complex flow structure is observed but also time-dependent supersonic flow visualization is realized. The mechanism of interaction between shock wave and turbulence is discussed based on high quality NPLS image.     

Experimental investigation of supersonic flow over a hemisphere

The experimental investigation of supersonic flow over a hemisphere was conducted using Nanoparticle-based Planar Laser Scattering(NPLS) technique in a supersonic quiet wind tunnel at Ma=2.68.Ahead of the hemisphere,boundary layer separation with the formation of a three-dimensional separated flow was observed,which was resulted from the interaction between the three-dimensional bow shock wave and the boundary layer.The complex flow structures of supersonic flow over the hemisphere were visualized.Based on the time correlation of NPLS images,time-space evolutionary characteristics of supersonic flow over the hemisphere were studied,and the evolutionary characteristics of the spanwise and streamwise large scale vortex structures were obtained,which have the features of periodicity and similar geometry.     

An experimental study on fine structures of supersonic laminar/turbulent flow over a backward-facing step based on NPLS

Fine structures of supersonic flow over a 5 mm high backward facing step(BFS),including expansion wave fan,reattachment shock,supersonic boundary layer were measured in a Ma=3.0 low-noise indraft wind tunnel.By varying the superficial roughness of the wall upstream from the step,supersonic laminar flow and supersonic turbulent flow could be formed over a BFS.Measurements on the spatiotemporal features of the holistic flow field and the fine structures in four typical regions were carried out using NPLS(nano-based planar laser scattering).Flow structures,including expansion wave fan,reattachment shock,supersonic boundary layer and its separation,reattachment and redevelopment are revealed by measuring the holistic structure of the transient flow field.Comparing the two time-averaged flow fields with each other,it is apparent that supersonic turbulent flow over a BFS(STF-BFS) has a larger expansion angle and a shorter recirculation region,and its redeveloped boundary layer increases at a smaller obliquity while the angle of reattachment shock is the same for the supersonic laminar flow over a BFS(SLF-BFS).With regard to time-evolution features,the K-H vortices in the SLF-BFS suffers from shearing,expansion,reattachment and three-dimensional effects while in the STF-BFS large-scale structures are affected by the incline and distortion at the reattachment point due to expansion,viscosity and reverse-pressure.Studies on local regions indicate that in the SLF-BFS,the emergence of compression waves which distinctly converge into a reattachment shock is due to the local convective Mach number and the inducement of K-H vortices in the free shear layer.Nevertheless,in the STF-BFS,compression waves and K-H vortices are barely evident,and the formation of a reattachment shock is related to the wall compressive effect.