具有局部稀薄气体效应的高超声速尖锥气动加热特征研究

以微钝尖锥为飞行器前缘模型,采用基于分子运动论的DSMC方法模拟不同前缘曲率半径的尖锥在高超声速来流下的气动热环境,计算驻点热流率,并与Fay-Riddell公式和其他修正理论作对比,研究具有局部稀薄气体效应的高超声速尖锥气动加热特征及其变化规律.发现修正的Cheng参数适合作为工程上判断驻点区域稀薄气体效应影响大小的判据.     

高超声速飞行器气动热环境的化学非平衡效应数值研究（英文）

针对高超声速飞行器气动热环境的化学非平衡效应进行数值模拟.首先以二维圆柱钝体为模型,采用三种气体模型比较和分析化学非平衡效应对气动特性和气动加热特性的影响;接着以三维球头钝体为模型,进一步分析化学非平衡效应及壁面催化特性对气动加热特性的影响.分析表明,化学非平衡效应对压力分布的影响不大;但它能显著缩短激波脱体距离,更重要的是,能大幅降低波后温度,对飞行器气动热环境产生深远影响.另外,壁面催化特性对传热也有较大的影响,对壁面材料催化进程的有效控制可以大幅降低壁面传热.     

前缘一体化高温热管结构防热效果的实验研究

为了研究复杂构型前缘一体化高温热管结构在高热流密度状态下的防热效果, 设计了飞行器气动加热轨道, 实现了高温热管低状态完全启动、高状态极限考核。然后采用超声速电弧风洞驻点自由射流结合轨道模拟技术, 模拟乘波体飞行器的前缘疏导构件的气动加热环境, 开展了前缘一体化高温热管结构防热效果研究。实验结果表明, 一体化高温热管结构能够多次使用, 低状态下高温热管的启动时间约为115 s, 在高状态下结构依然有效, 降温系数达到24.5%, 验证了前缘疏导式防热结构的防热效果, 可为未来新型高超声速飞行器非烧蚀热防护系统的设计提供指导。      

高超声速飞行器热管冷却前缘结构一体化建模分析

针对高超声速飞行器工作时前缘恶劣的气动加热环境, 为了保证飞行器前缘的尖锐外形, 提出内嵌高温热管前缘结构. 针对热管内部复杂流动与换热情况, 对内嵌高温热管前缘结构进行一体化建模, 将模型的核心部件液态金属热管工作状况的计算与实验进行对比以验证模型的可靠性. 本文还分析了给定工况下内嵌高温热管前缘结构的热防护效果, 其中壁面最高温度下降了11.6%, 最低温度上升了8%, 高温区和低温区均被封闭在前缘外层区域, 内层温度更加均匀, 实现了热流由高温区向低温区的转移, 削弱了高温区的热负荷. 本文还分析了接触热阻对热管冷却前缘结构效果的影响. 关键词： 热管 前缘 疏导式热防护 气动热     

火星进入热环境预测的热力学模型数值分析

采用数值模拟目的 研究不同热力学模型中火星飞行器表面的热环境.研究发现,驻点区域流动近似为热力学平衡,随着飞行高度增加,流动逐渐偏向热力学非平衡.当壁面为非催化壁时,不同热力学模型所得热流基本一致,当壁面为完全催化壁时,热力学非平衡模型所得热流更高,这种差异由化学反应特征温度差异引起,并随着流动的热力学非平衡特性增加而增加.     

高超声速飞行器前缘疏导式热防护结构的实验研究

针对高超声速飞行器前缘疏导式防热结构的特点,设计前缘内嵌高导热率材料结构和一体化层板热管结构两类对比实验,用于验证前缘疏导式防热结构的可行性.利用球形短弧氙灯作为辐射热源模拟气动加热,分别对钢质前缘、内嵌铜材料的钢质前缘和一体化层板式热管前缘进行加热,测量前缘驻点区域和尾部翼面区域的温度变化.实验结果表明:内嵌高导热率材料的前缘疏导结构能够降低头部驻点区的温度,提高尾部低温区的温度,实现对前缘结构的热防护;以蒸馏水作为工质一体化层板式热管前缘结构,在较低热流条件下也能够实现对前缘驻点区的疏导式热防护,但在较高热流条件下,由于水蒸气压力过大使得层板式前缘结构发生破坏,体现出热管内部工作介质对结构防热效果和应用范围都起到的关键作用.     

各向异性导热率对前缘结构热防护性能的影响

高超声速飞行器在高马赫数飞行时,前缘结构要经历严酷的气动加热。炭/炭复合材料是用于前缘结构的理想材料,由于炭/炭复合材料具有各向异性和高导热率,因此本文研究了材料热导率和各向异性结构,以及涂层的热导率对驻点温度的影响。数值模拟的结果表明,基体材料沿热传导方向的热导率对驻点温度有显著的影响,其它方向的热导率对驻点温度影响不大。在基体材料热导率较低的情况下,涂层对驻点温度的影响明显,当基体材料的热导率较高时,涂层对驻点温度的影响较小。     

二维高超声速后台阶表面传热特性实验研究

高超声速后台阶流动是大气层内高速飞行器发动机设计、表面热防护以及高超声速拦截器红外成像窗口气动光学效应校正等诸多先进高超声速技术研发过程中所涉及的一类基础流动问题. 研究高超声速后台阶流动特性对有效提升飞行器综合性能, 进一步掌握高超声速流动机理具有重大基础 意义. 本文以二维高超声速后台阶流动为研究对象, 在KD-01高超声速激波风洞中测量了二维后台阶模型表面传热系数和表面静压, 并将实测台阶下游表面传热系数分布同采用高超声速边界层理论所得估计值进行了比较. 为进一步验证实验结果, 使用NPLS技术测量了其中一种实验状态下台阶周围流动结构. 研究发现, 对于二维高超声速后台阶流动, 台阶下游表面传热分布受台阶处边界层外缘流动特性的直接影响; 在台阶下游分离区和再附区内, 气体黏性占主导作用; 在台阶下游远场区域, 边界层流动特性趋同于平板边界层; 下游边界层基本结构取决于台阶处边界层相对厚度. 对高超声速后台阶流动, 若使用数值模拟方法研究气动热问题, 应当使用湍流模型.     

无限薄平板边界层前缘感受性过程的数值研究

在边界层流中层流向湍流转捩机理的研究一直是人们所关注的重要理论课题之一.感受性阶段是边界层内整个转捩过程中的初始阶段,它在层流向湍流转捩过程中起着关键性的作用.但是,在过去的边界层前缘感受性研究中大多数都是针对外部声波小扰动,很少见到考虑在自由流中普遍存在的自由来流湍流作用下边界层内诱导前缘感受性问题的相关报道.本文采用直接数值模拟的方法,研究自由来流湍流与无限薄平板前缘驻点扰动作用下边界层流中前缘感受性过程的内在机理.数值结果发现,在自由来流湍流与无限薄平板前缘驻点扰动作用下边界层流中能被感受出一组小扰动波,且它们的色散关系、增长率、中性曲线等结果都与流动稳定性中的线性理论获得的理论解相一致,由此可知在边界层内被激发产生的一组小扰动波就是Tollmien-Schlichting波,这也证明自由来流湍流与无限薄平板前缘驻点扰动相互作用是激发边界层流中前缘感受性过程的另一种物理机理;另外,还探讨了自由来流湍流度以及自由来流湍流的运动方向对无限薄平板边界层前缘感受性过程有何影响等.总之,开展边界层前缘感受性过程的深入研究,有益于完善流动稳定性理论,将为层流向湍流转捩过程的预测提供合理的理论依据.     

吸气式高超声速飞行器气动力气动热的数值模拟方法及应用

对吸气式高超声速飞行器而言,物面热流和摩阻的准确预测对飞行器设计及安全十分关键.介绍采用CFD准确预测气动力和气动热的方法,包括流动的控制方程、湍流模型及湍流的先进壁面函数边界条件,介绍流动的数值求解方法.对典型超声速层流和湍流流动的摩擦阻力和热流进行详细的验证与确认,考察CFD工具在使用先进壁面函数边界条件后,湍流计算的法向网格无关性能力.对设计的一种吸气式高超声速飞行器的气动力和气动热进行数值模拟,为飞行器的气动设计及热防护提供了可靠的数据.     

Variation character of stagnation point heat flux for hypersonic pointed bodies from continuum to rarefied flow states and its bridge function study

This paper is a research on the variation character of stagnation point heat flux for hypersonic pointed bodies from continuum to rarefied flow states by using theoretical analysis and numerical simulation methods. The newly developed near space hypersonic cruise vehicles have sharp noses and wingtips, which desires exact and relatively simple methods to estimate the stagnation point heat flux. With the decrease of the curvature radius of the leading edge, the flow becomes rarefied gradually, and viscous interaction effects and rarefied gas effects come forth successively, which results in that the classical Fay-Riddell equation under continuum hypothesis will become invalid and the variation of stagnation point heat flux is characterized by a new trend. The heat flux approaches the free molecular flow limit instead of an infinite value when the curvature radius of the leading edge tends to 0. The physical mechanism behind this phenomenon remains in need of theoretical study. Firstly, due to the fact that the whole flow regime can be described by Boltzmann equation, the continuum and rarefied flow are analyzed under a uniform framework. A relationship is established between the molecular collision insufficiency in rarefied flow and the failure of Fourier’s heat conduction law along with the increasing significance of the nonlinear heat flux. Then based on an inspiration drew from Burnett approximation, control factors are grasped and a specific heat flux expression containing the nonlinear term is designed in the stagnation region of hypersonic leading edge. Together with flow pattern analysis, the ratio of nonlinear to linear heat flux W _r is theoretically obtained as a parameter which reflects the influence of nonlinear factors, i.e. a criterion to classify the hypersonic rarefied flows. Ultimately, based on the characteristic parameter W _r , a bridge function with physical background is constructed, which predicts comparative reasonable results in coincidence well with DSMC and experimental data in the whole flow regime.     

Aerothermodynamic Analysis of a Reentry Brazilian Satellite

This work deals with a computational investigation on the small ballistic reentry Brazilian vehicle SAtélite de Reentrada Atmosférica (SARA). Hypersonic flows over the vehicle SARA at zero-degree angle of attack in chemical equilibrium and thermal nonequilibrium are modeled by the direct simulation Monte Carlo method, which has become the main technique for studying complex multidimensional rarefied flows, and which properly accounts for the nonequilibrium aspects of the flows. The emphasis of this paper is to examine the behavior of the primary properties during the high-altitude portion of SARA reentry. In this way, velocity, density, pressure, and temperature field are investigated for altitudes of 100, 95, 90, 85, and 80?km. In addition, comparisons based on geometry are made between axisymmetric and planar two-dimensional configurations. Some significant differences between these configurations were noted on the flowfield structure in the reentry trajectory. The analysis showed that the flow disturbances have different influence on velocity, density, pressure, and temperature along the stagnation streamline ahead of the capsule nose. It was found that the stagnation region is a thermally stressed zone. It was also found that the stagnation region is a zone of strong compression, high wall pressure. Wall pressure distributions are compared with those of available experimental data, and good agreement is found along the spherical nose for the altitude range investigated.     

材料防热的多尺度现象与防热建模

新型高速飞行器返回再入及在大气层内飞行的过程中面临多样化的气动加热环境,材料工艺的改进和多组分的添加使材料高温热响应特性变得更加复杂,并呈现多尺度特性.文章从飞行热环境、材料工艺特征和细微观响应等方面对材料防热机理和建模方法进行了阐述,对不同类型飞行器热环境特征与防热建模难点进行了分析,对各类防热材料工艺与热响应特点进...     

跨流域高超声速绕流环境Boltzmann模型方程统一算法研究

如何准确可靠地模拟从外层空间高稀薄流到近地面连续流的航天器高超声速绕流环境与复杂流动变化机理是流体物理的前沿基础科学问题. 基于对Boltzmann方程碰撞积分的物理分析与可计算建模, 确立了可描述自由分子流到连续流区各流域不同马赫数复杂流动输运现象统一的Boltzmann模型速度分布函数方程, 发展了适于高、低不同马赫数绕流问题的离散速度坐标法和直接求解分子速度分布函数演化更新的气体动理论数值格式, 建立了模拟复杂飞行器跨流域高超声速飞行热环境绕流问题的气体动理论统一算法. 对稀薄流到连续流不同Knudsen数0.002 ≤Kn_∞ ≤1.618、不同马赫数下可重复使用卫星体再入过程(110–70 km)中高超声速绕流问题进行算法验证分析, 计算结果与典型文献的Monte Carlo直接模拟值及相关理论分析符合得较好. 研究揭示了飞行器跨流域不同高度高超声速复杂流动机理、绕流现象与气动力/热变化规律, 提出了一个通过数值求解介观Boltzmann模型方程, 可靠模拟高稀薄自由分子流到连续流跨流域高超声速气动力/热绕流特性统一算法.     

Multi-scale k-distribution model for gas mixtures in hypersonic nonequilibrium flows

A k-distribution model is presented for gas mixtures in thermodynamic nonequilibrium, containing strongly radiating atomic species N and O together with molecular species of N₂, N₂⁺, NO and O₂. In the VUV range of the spectrum there is strong absorption of atomic radiation by bands of N₂. For this spectral range, a multi-scale model is presented, where RTEs are solved separately for each emitting species and overlap with other species is treated in an approximate way. Methodology for splitting the gas mixture into scales and evaluation of the overlap parameter between different scales is presented. The accuracy of the new model is demonstrated by solving the radiative transfer equation along the stagnation line flow field of the Crew Exploration Vehicle (CEV).     

稀薄气体流动非线性耦合本构关系研究进展

临近空间高超声速飞行器流场蕴含着复杂的非线性流动机理与丰富的热化学非平衡流动现象, 基于Newton摩擦定律和Fourier热传导定律的Navier-Stokes(N-S)方程不足以描述高超声速飞行器从连续流到稀薄流的多尺度非平衡现象。非线性耦合本构关系(nonlinear coupled constitutive relations, NCCR)作为一种全新的本构方程体系, 在严格满足热力学熵条件的基础上, 巧妙地构建了应力与热流的非线性表达形式。然而, NCCR方程的强非线性耦合特性是求解过程的一大难题。为了克服这一技术瓶颈, 提出了混合迭代算法, 为实现NCCR方程的高效稳定求解提供了坚实的理论基础。在该理论研究的基础上, 考虑到原始NCCR方程对热通量演化方程的简化处理, 降低了方程的计算精度, 提出了改进的NCCR+方程。该方程在强激波压缩区域和膨胀区域表现出比传统NCCR方程更高的计算精度与更强的非平衡流动模拟能力。同时, 为了解决临近空间高超声速空气动力学的多尺度与多物理效应耦合难题, 提出了NCCR与转动非平衡的耦合计算模型, 拓展了NCCR方程在双原子气体中的模拟能力。为了揭示稀薄气体效应与真实气体效应的耦合作用机理, 进一步建立了NCCR与热化学反应的耦合计算方法。大量研究结果表明, 考虑多物理效应的NCCR方程在低Kn下能够恢复到与N-S方程一致的解。随着Kn的增加, 流场的非平衡程度逐渐增强, 其结果与N-S方程差异显著, 而与DSMC方法计算结果和实验数据具有更好的一致性。      

气动加热对高超声速飞行器激光毁伤效应影响

通过采用工程计算方法求解高超声速飞行器碳-碳复合材料分别在气动热、激光单独作用以及气动热/激光耦合作用下的热化学烧蚀。计算分析表明:激光单独作用下,碳-碳复合材料的烧蚀速率较小;随激光能量的增加,碳-碳复合材料的烧蚀速率增加;气动加热条件下激光对高超声速飞行器碳-碳复合材料的烧蚀毁伤效应会明显增强;沿弹道的气动加热累积效应对碳-碳复合材料气动热/激光耦合烧蚀作用不明显。     

A hybrid particle approach for continuum and rarefied flow simulation

A hybrid particle scheme is presented for the simulation of compressible gas flows involving both continuum regions and rarefied regions with strong translational nonequilibrium. The direct simulation Monte Carlo (DSMC) method is applied in rarefied regions, while remaining portions of the flowfield are simulated using a DSMC-based low diffusion particle method for inviscid flow simulation. The hybrid scheme is suitable for either steady state or unsteady flow problems, and can simulate gas mixtures comprising an arbitrary number of species. Numerical procedures are described for strongly coupled two-way information transfer between continuum and rarefied regions, and additional procedures are outlined for the determination of continuum breakdown. The hybrid scheme is evaluated through a comparison with DSMC simulation results for a Mach 6 flow of N₂ over a cylinder, and good overall agreement is observed. Large potential efficiency gains (over three orders of magnitude) are estimated for the hybrid algorithm relative to DSMC in a simple example involving a rarefied expansion flow through a small nozzle into a vacuum chamber.