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PIV技术在超及高超声速流场测量中的研究进展 总被引:1,自引:0,他引:1
本文分析了超声速流场对测量技术的特殊要求, 归纳了目前将粒子影像测速仪(particle image velocimetry, PIV)技术应用于超声速流场的测量时所面临的主要技术难点以及主要的解决方法, 分析了超声速流场中所用PIV粒子的主要要求、粒子特性、投放方法等, 介绍了PIV技术在超声速、高超声速流场测量中最新的国内外进展, 特别是给出了国内外关于高超声速流场中激波/附面层的相互干扰, 以及高超声速飞行器超燃冲压发动机主要部件内流场的PIV试验研究的最新进展. 相似文献
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为研究活塞回复运动对火药燃气流动的影响,基于两相流理论对活塞控制侧向后喷武器的发射过程进行了数值模拟研究。考虑控制侧向后喷通道开闭的活塞-弹簧系统的往复运动,建立了结合膛内气固两相流、活塞腔内流固耦合和侧向排气管内气体瞬态流动的武器发射过程数学模型,并将数值模拟结果与相关文献进行了比较验证。得到了该武器发射过程中膛内流场分布与稀疏波传播特性,并与普通武器的膛内流场进行了对比分析。进一步研究了活塞回复运动对火药燃气流动和减后坐效率的影响。结果表明:相对于不考虑活塞的回复运动,在弹丸初速都降低1.52%的情况下,因为活塞回复关闭后喷通道,其减后坐效率由38.86%下降到32.88%,说明在此类武器研究中,不可忽视活塞回复运动。 相似文献
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本文采用分区搭接网格技术,对机翼/机身/挂架/短舱复杂组合体进行网格分布,通过分析计算网格对结果的影响,探讨了网格的划分.基于Roe的近似黎曼解的方法,采用S-A湍流模型,通过求解N-S方程,对该组合体外流场/发动机短舱内流场进行一体化数值模拟,与相应风洞实验数据进行了比较与分析,取得了与实验数据较为吻合的结果.与无发动机短舱的组合体的气动特性进行比较,分析了短舱对翼身组合体的气动干扰. 相似文献
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采用通量修正二阶TVD格式和混合隐式时间推进/解析算法求解k-ε湍流模型的输运方程以克服其计算刚性. 通过过膨胀状态下轴对称喷管内流场、带圆转方型面二元喷管内外流场两个算例, 验证了其配合改进的多重网格算法对计算效率的提高. 计算结果与试验数据进行了对比. 相似文献
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载人列车车厢内空气流场温度场数值模拟 总被引:4,自引:0,他引:4
采用稳态不可压缩雷诺时均N-S方程、k-ε湍流模型,计算了载人列车车厢内三维空气流场和温度场。将太阳辐射热和人体散热作为能量方程的附加源项,研究了在条缝送风条件下,乘客和太阳辐射对车厢内流场和温度场的影响。计算结果表明:现有的送风方式除车厢两端外,车厢内沿长度方向气流分布比较均匀;送风口的布置对车厢内流场温度场分布影响较大,送风气流在车厢两侧形成两股比较大的流动旋涡;人体散热和太阳辐射对车厢内流场温度场影响较大,非空载时车厢内流场分布与空载时有较大差别,太阳照射和人体产生的热气流使车厢内存在较大的温度梯度,靠窗处的温度较高,过道处温度较低。流场温度场的计算结果和实验数据吻合较好。 相似文献
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拼接网格技术在复杂流场数值模拟中的应用研究 总被引:1,自引:0,他引:1
采用分区拼接网格技术,对 DLR-F6 机翼/机身胜架/短舱复杂组合体进行拼接网格分布.并采用 Menter SST 湍流模型,通过求解 Navier-Stokes 方程,对该组合体外流场以及发动机短舱内流场进行了一体化数值模拟,与相应风洞实验数据及分区搭接网格计算结果进行了比较与分析,验证了拼接网格技术的高效性与可靠性.同时通过分析对比不同插值方法的计算结果,研究了插值方法对拼接精度的影响;通过分析对比几组不同的拼接网格算例,总结出了 3 个拼接网格的基本实施准则.证明了拼接网格能够大幅度减小计算网格数目,可以更加灵活地分布网格节点,这样既可以缩短计算时间,又可以降低对内存的需求,提高了计算效率;同时无论整体的力系数,还是局部的压力分布流场细节都能够满足工程精度. 相似文献
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超声速气流磁流体加速初步实验研究 总被引:5,自引:2,他引:3
利用激波风洞, 采用氦气驱动氩气, 在平衡接触面运行方式下得到高温气体,通过在低压段注入电离种子K2CO3粉末, 实现高温条件下导电流体的产生, 设计了超声速喷管及磁流体加速实验段, 采用大电容提供电能, 开展了超声速气流磁流体加速初步实验研究. 典型实验条件下, 当喷管入口总压为0.7049MPa、理论平衡温度为8372.8K, 喷管出口马赫数为1.5, 电容充电电压为400V, 磁感应强度为0.5T时, 对电压电流特性、电导率、负载系数、电效率、加速效果等进行了测量或计算, 主要结论有: 磁场作用下的超声速气流的电导率的值大约在150S/m; 磁流体加速通道负载系数约为4, 电效率约为28%, 平均输入功率约198kW; 采用电参数测试方法对磁流体加速效果进行评估, 速度增加约15.7%;超声速气流的电导率对加速通道的电效率及加速效果等有很重要的影响. 相似文献
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高超声速飞行器动力系统研究进展 总被引:20,自引:0,他引:20
简要介绍了高超声速飞行器动力系统的概况.第2部分介绍了超燃冲压发动机、爆震发动机和组合循环发动机等典型高超声速吸气式发动机的基本工作原理与系统组成,描述了各自的特点.第3部分阐述了高超声速飞行器动力系统存在的难点问题,并列出了在总体设计、进气道、燃烧室、尾喷管、热防护、轻质结构、燃油供应与控制等方面的关键技术.第4部分回顾了上述几种典型发动机的发展历程,比较全面地介绍了世界主要航空、航天大国在动力系统关键技术攻关与系统研制方面的主要研究计划和取得的主要进展,总结了经验教训, 指出了发展趋势.第5部分阐述了高超声速飞行器动力系统中的燃烧过程及其燃烧基本问题,介绍了主要研究进展. 相似文献
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Yu. V. Tunik 《Fluid Dynamics》2011,46(5):775-781
Three variants of the startup of an axisymmetric convergent-divergent nozzle are considered with the static pressures at the
entry and exit of the nozzle being the same at the beginning of the process. The subsonic startup corresponds to open nozzle
acceleration in air. The supersonic startup simulates the sudden opening of a cover at the nozzle inlet under supersonic flight
conditions. A successful nozzle startup with the formation of steady supersonic flow along the whole channel is realized in
the third variant of supersonic startup with gas injection through a small region of the wall of the divergent nozzle section.
The investigation is performed numerically, on the basis of the Euler equations for axisymmetric gas flows. 相似文献
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We establish the existence and stability of multidimensional steady transonic flows with transonic shocks through an infinite
nozzle of arbitrary cross-sections, including a slowly varying de Laval nozzle. The transonic flow is governed by the inviscid
potential flow equation with supersonic upstream flow at the entrance, uniform subsonic downstream flow at the exit at infinity,
and the slip boundary condition on the nozzle boundary. Our results indicate that, if the supersonic upstream flow at the
entrance is sufficiently close to a uniform flow, there exists a solution that consists of a C
1,α subsonic flow in the unbounded downstream region, converging to a uniform velocity state at infinity, and a C
1,α multidimensional transonic shock separating the subsonic flow from the supersonic upstream flow; the uniform velocity state
at the exit at infinity in the downstream direction is uniquely determined by the supersonic upstream flow; and the shock
is orthogonal to the nozzle boundary at every point of their intersection. In order to construct such a transonic flow, we
reformulate the multidimensional transonic nozzle problem into a free boundary problem for the subsonic phase, in which the
equation is elliptic and the free boundary is a transonic shock. The free boundary conditions are determined by the Rankine–Hugoniot
conditions along the shock. We further develop a nonlinear iteration approach and employ its advantages to deal with such
a free boundary problem in the unbounded domain. We also prove that the transonic flow with a transonic shock is unique and
stable with respect to the nozzle boundary and the smooth supersonic upstream flow at the entrance. 相似文献
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Hisashi Mikami 《国际流体数值方法杂志》1987,7(6):603-619
The piecewise linear method (PLM) based on time operator splitting is used to solve the unsteady compressible Euler equations describing the two-dimensional flow around and through a straight wall inlet placed stationary in a rapidly rotating supersonic flow. The PLM scheme is formulated as a Lagrangian step followed by an Eulerian remap. The inhomogeneous terms in the Euler equations written in cylindrical coordinates are first removed by Sod's method and the resulting set of equations is further reduced to two sets of one-dimensional Lagrangian equations, using time operator splitting. The numerically generated flow fields are presented for different values of the back pressure imposed at the downstream exit of the inlet nozzle. An oblique shock wave is formed in front of the almost whole portion of the inlet entrance, the incoming streamlines being deflected towards the higher pressure side after passing through the oblique shock wave and then bending down to the lower pressure side. A reverse flow appears inside the inlet nozzle owing to the recovery pressure of the incoming streams being lower than the back pressure of the inlet nozzle. 相似文献
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关于吸气式高超声速推进技术研究的思考 总被引:5,自引:0,他引:5
回顾了吸气式高超声速推进技术的研究进展, 分析了超燃冲压发动机研制面临的关键科学问题, 并从不同角度探讨了增大超燃冲压发动机推力的可能方法.这些方法包括: 能够降低总压损失的高超声速来流压缩方法、生成三维涡流的超声速混合增强技术、碳氢燃料的预热喷射、可以控制燃烧过程的燃烧室设计优化方法、通过减小发动机流道湿面积来降低摩擦阻力和催化复合解离的燃气降低高温气体效应.考虑到等压热力学循环的热效率,还建议研究在高超声速推进系统中应用热效率高的爆轰过程, 并探讨了爆轰推进方法研究的进展与问题.吸气式高超声速推进技术是高超声速飞行器发展的关键技术, 认真思考和探索其发展方向是非常必要的. 相似文献
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Experimental investigation of combustion mechanisms of kerosene-fueled scramjet engines with double-cavity flameholders 总被引:2,自引:0,他引:2
Yu Pan Jian-Guo Tan Jian-Han Liang Wei-Dong Liu Zhen-Guo Wang 《Acta Mechanica Sinica》2011,27(6):891-897
A scramjet combustor with double cavitybased flameholders was experimentally studied in a directconnected test bed with the inflow conditions of M = 2.64,Pt = 1.84 MPa,Tt = 1 300 K.Successful ignition and selfsustained combustion with room temperature kerosene was achieved using pilot hydrogen,and kerosene was vertically injected into the combustor through 4×φ 0.5 mm holes mounted on the wall.For different equivalence ratios and different injection schemes with both tandem cavities and parallel cavities,flow fields were obtained and compared using a high speed camera and a Schlieren system.Results revealed that the combustor inside the flow field was greatly influenced by the cavity installation scheme,cavities in tandem easily to form a single side flame distribution,and cavities in parallel are more likely to form a joint flame,forming a choked combustion mode.The supersonic combustion flame was a kind of diffusion flame and there were two kinds of combustion modes.In the unchoked combustion mode,both subsonic and supersonic combustion regions existed.While in the choked mode,the combustion region was fully subsonic with strong shock propagating upstream.Results also showed that there was a balance point between the boundary separation and shock enhanced combustion,depending on the intensity of heat release. 相似文献
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In the framework of the European Commission co-funded LAPCAT (Long-Term Advanced Propulsion Concepts and Technologies) project, the methodology of a combined ground-based testing and numerical modelling analysis of supersonic combustion flow paths was established. The approach is based on free jet testing of complete supersonic combustion ramjet (scramjet) configurations consisting of intake, combustor and nozzle in the High Enthalpy Shock Tunnel Göttingen (HEG) of the German Aerospace Center (DLR) and computational fluid dynamics studies utilising the DLR TAU code. The capability of the established methodology is demonstrated by applying it to the flow path of the generic HyShot II scramjet flight experiment configuration. 相似文献