变截面主流加速对超音速气膜冷却的影响

本文对扩张通道和直通道中的超音速气膜冷却进行了数值模拟,分析了主流加速和不加速两种情况下超音速气膜冷却的差异.计算结果表明,在主流加速的情况下,超音速气膜冷却的效果要好于相同进口条件下主流不加速的情况,并且这种差异越到下游越明显.同时,在本文计算的工况范围中,无论主流加速还是不加速,增大主流进口马赫数或者减少冷却流进口马赫数都使气膜冷却效率下降,而壁面静压则由主流气体决定.     

高超声速平板边界层流动显示的试验研究

本文在高超声速脉冲式风洞内对基于纳米示踪的平面激光散射技术(nano-based planar laser scattering, NPLS)的应用进行了探索, 并在此基础上对平板边界层流动结构的精细测量进行了研究. 试验来流Ma=7.3, 总压4.8 MPa, 总温680 K. 通过时序的分析和调试, 对各分系统实现了高精度的同步控制; 定量的粒子注入及混合, 实现了粒子的均匀撒播, 对主流获得了均匀的显示效果; 对于边界层流动, 获得了精细的瞬态流动结构图像, 显示了层流到湍流的转捩过程, 并分析了其时空演化特性.     

Variation of the temperature factor of models on a ballistic range

A technique for variation of the temperature factor of free-flight models by varying their initial temperature is described. An experiment on a ballistic range is carried out with a free-flying supersonic blunt cone with a half-angle of 15° at a Mach number of 2.3. The flow at the cone base is studied in the transition range (from the laminar to turbulent flow) of Reynolds numbers. The base flow pattern is determined from the shadowgraphs of the flow about the models. The drag coefficient of the blunt cone at a zero angle of attack is found by processing trajectory data. It is found that the near wake geometries and the drag coefficients of the models tested at the laboratory temperature and a temperature of 120 K differ. Explanations of this effect are given.     

Measurement and simulation of edge plasma flow in the Tore Supra tokamak

A simple fluid model is used to interpret two edge flow phenomena that are observed in the scrape-off layer of the Tore Supra tokamak. (1) The Mach number of the parallel flow midway between the two sides of the toroidal limiter is always a significant fraction of the ion sound speed, typicallyM _?=?0.3÷?0.6. directed towards the high field side. In Tore Supra theE×B drift is relatively unimportant, so the esixtence of alarge parallel flow implies a significant poloidal aysmmetry of the SOL source. (2) Massive injection of neutral gas leads to a recycling source that is poloidally localized near the nozzle. The parallel Mach number changes during injection. The sign of the change is positive for injection on the high field side, and negative for injection on the low field side. The ionization source that is needed to produce the change is about 5÷10% of the total amount of injected gas, consistent with atomic physics calculations.     

Tm3+掺杂玻璃激光致冷新方案的理论分析

提出了一种用于块状Tm3 掺杂玻璃ZBLANP(ZrF4-BaF2-LaF3-NaF-PbF2)激光致冷的新方案,即通过两组相互垂直的高反射率平面镜的多次反射,增加对抽运光的吸收,为了减少抽运光在样品表面入射时的反射损失,抽运光以布儒斯特角入射。重点就该方案的样品致冷温度和时间常量进行了详细的理论计算与分析,从理论上分析了入射激光功率为1 W时致冷功率和反射次数的关系,研究了不同入射激光功率下致冷温度和反射次数的关系。结果表明,入射激光功率分别为3 W、4 W5、W时,样品可以分别从室温(300 K)被冷却到275.7 K、267.4 K、259.1 K。特别是在入射激光功率为4.5 W时,样品从室温开始被冷却到263.2 K,即致冷后样品的温度下降了36.8K,比传统致冷方法提高了约53.3%,相应的致冷时间常量为22.7 min。     

Combined energy–power action on a source in the constant Mach number regime with the given external force

Combined action on a source that flows into a submerged area or vacuum in the constant Mach number regime has been studied. The action by an external force has been defined with a constant distribution function (the force is given per unit volume) and with a distribution function proportional to the gas density (the force is given per unit mass). The investigations have been carried out for cylindrical and spherical sources. Similarity and differences, advantages and drawbacks of the above-mentioned cases and variants have been analyzed. It has been shown that the enthalpy increases significantly in subsonic flow (for the Mach number smaller than unity) by several times in the cylindrical source and by more than an order of magnitude in the spherical source. The total enthalpy increment increases with the length of the action zone or with the coordinate of the closing section.     

FD-21风洞Ma=10高超声速推进试验技术探索

针对Mach数8以上（Ma>8）冲压发动机地面试验能力不足问题,基于FD-21高能脉冲风洞,开展了吸气式推进试验技术探索,提升了FD-21风洞的重活塞驱动能力,获得了总压18.66 MPa、总温3 950 K、Ma=9.62、静压436.6 Pa、速度3 km/s的高焓大动压模拟流场,同时发展了高时间分辨率吸收光谱测量技术和基于重模型自由飞原理的发动机推阻测量方法.在此基础上,设计了弯曲激波压缩二元发动机,构建了燃料在线供应与喷注控制、模型悬挂与瞬态释放及相关测量一体的试验系统,在所建立的Ma=9.62风洞模拟环境中进行了集成验证试验,定量测得了有/无氢气射流与空气/氮气超声速气流作用下二元发动机的壁面压力、吸收光谱峰值吸收率、轴向力等数据,并利用纹影观测到了进气道唇口与燃烧室部位的波系特征.多次试验所得的壁面压力、峰值吸收率、轴向力随时间变化曲线均存在2 ms以上的平台,表明二元发动机建立了准定常流动.冷热态及氮气对照组对应的壁面压力分布、峰值吸收率、轴向力等数据呈现出了明显不同,且二者规律近似一致,一方面说明所建立的模拟流场、燃烧诊断技术、发动机推阻测量技术是有效的,另一方面也表明二元发动机实现了点火燃烧、获得有效热功转换,为后续相关研究奠定了良好的基础.      

新建高焓激波风洞Ma=8飞行模拟条件的实现与超燃实验

针对高Mach数超燃冲压发动机实验能力空缺问题,基于航天十一院新建的FD-21高能脉冲风洞,进行了Ma=8超燃飞行条件的模拟能力设计与调试,获得了总焓2.9 MJ/kg、总压11.01 MPa实验条件,实现了Ma=8、高度31 km飞行条件的风洞模拟.在此基础上,研发了匹配的氢燃料供应及喷注时序控制系统,设计了超燃冲压发动机模型,开展了超燃冲压发动机模型自由射流应用性风洞实验,获得了氢气燃料与空气、氮气超声速气流耦合流动作用下的实验模型壁面压力数据.在当量比近似一致条件下,空气来流对应的燃烧室壁面压力明显高于氮气来流情况,表明氢气在1 ms有效实验时间内完成了与超声速空气来流的混合、点火与燃烧,获得燃烧释热特性,确认了在FD-21高能脉冲风洞开展高Mach数超燃实验是切实可行的,为后续研究奠定了良好的基础.      

Effects of variable specific heat on energy transfer in a high-temperature supersonic channel flow

An energy transfer mechanism in high-temperature supersonic turbulent flow for variable specific heat (VSH) condition through turbulent kinetic energy (TKE), mean kinetic energy (MKE), turbulent internal energy (TIE) and mean internal energy (MIE) is proposed. The similarities of energy budgets between VSH and constant specific heat (CSH) conditions are investigated by introducing a vibrational energy excited degree and considering the effects of fluctuating specific heat. Direct numerical simulation (DNS) of temporally evolving high-temperature supersonic turbulent channel flow is conducted at Mach number 3.0 and Reynolds number 4800 combined with a constant dimensional wall temperature 1192.60 K for VSH and CSH conditions to validate the proposed energy transfer mechanism. The differences between the terms in the two kinetic energy budgets for VSH and CSH conditions are small; however, the magnitude of molecular diffusion term for VSH condition is significantly smaller than that for CSH condition. The non-negligible energy transfer is obtained after neglecting several small terms of diffusion, dissipation and compressibility related. The non-negligible energy transfer involving TIE includes three processes, in which energy can be gained from TKE and MIE and lost to MIE. The same non-negligible energy transfer through TKE, MKE and MIE is observed for both the conditions.     

带喷流超声速后台阶流场精细结构及其运动特性研究

采用基于纳米示踪的平面激光散射技术(NPLS)对带超声速喷流的后台阶流动精细结构进行了研究. 来流马赫数为3.4, 喷流实测马赫数为2.45, 而名义马赫数为2.5. 结果清晰地揭示了激波、剪切层、混合层、Kelvin-Helmholtz涡、羊角涡及湍流大尺度结构等大量典型流场结构. 基于大量流场精细结构图像, 对典型位置处的结构进行了空间两点相关性分析, 在喷流混合层前端涡结构小于湍流充分发展的尾端, 结构角相对也小. 喷流工作时, 模型台阶下游表面由一薄层气膜覆盖. 获得了模型流向和不同高度展向平面内的流场结构, 对照纹影试验结果, 分析了流动特点及时间演化规律. 采用微型压力扫描系统测试了模型表面的压力系数分布, 靠近喷流下游处压力系数区域0.0146. 针对NPLS图像做了流动的分形维数的分析, 发现在流动初始阶段分形维数接近于1, 越靠下游分形维数越高.     

Heat transfer in lattice BGK modeled fluid

The thermal lattice BGK model is a recently suggested numerical tool aiming at solving problems of thermohydrodynamics. The quality of the lattice BGK simulation is checked in this paper by calculating temperature profiles in the Couette flow under different Eckert and Mach numbers. A revised lower order model is proposed to improve the accuracy and the higher order model is proved to be advantageous in this respect, especially in the flow regime with a higher Mach number.     

Effect of flow on quasi-one-dimensional acoustic wave propagation in a variable area duct of finite length

The general equation for the velocity potential of quasi-one-dimensional acoustic wave motion in a variable area, finite duct with one-dimensional flow is derived by using a perturbation technique. The non-linear second-order partial differential equation is linearized and then solved, by either a power series expansion method or the Runge-Kutta fourth-order method, for harmonic time dependence. The boundary condition taken at the duct mouth is that of matching the impedance of the duct sound field to that of the radiation field at the duct opening. Three axial Mach number variations along the duct axis are considered and the results obtained are compared with those for the case of constant Mach number, to determine the influence of the axial velocity gradient on sound propagation. The effect of flow on the radiation impedance is also considered.     

Effects of fuel properties on self-ignition and flame-holding performances in SCRAM jet combustor for n-alkane fuels

Combustion characteristics of liquid hydrocarbon fuels are studied in a model combustor of SCRAM jet engines. The Mach number and total pressure of main flow in the combustor are 2.0 and 0.38 MPa, respectively, and the total temperature is varied from 1800 to 2400 K. Five kinds of n-alkane fuels such as n-heptane, n-octane, n-decane, n-tridecane and n-hexadecane are employed in experiments. Fuels are injected with a carrier nitrogen gas perpendicular to the mail flow in the combustor and the self-ignition behavior is investigated. The results show that the liquid fuels with lower carbon number have better self-ignition performance. This suggests that physical properties of liquid fuels such as volatility have a dominant effect on the self-ignition. The flame-holding behavior is investigated with the addition of pilot hydrogen to carrier nitrogen gas. The critical equivalence ratio at which the stable combustion keeps after cut-off of the pilot hydrogen is obtained. The relationship between the critical equivalence ratio and carbon number of fuel shows that fuels with the carbon numbers from 8 to 10 have the best flame-holding performance among the tested fuels. These experimental results can be expressed qualitatively by the simplified analysis with the concept of physical and chemical induction times.     

压缩湍流剪切层-空腔流体气动光学效应的计算机仿真

对亚音速机载光电系统由于剪切层-腔流体而导致的气动光学效应进行了数值模拟.剪切层内的压缩效应用对流马赫数0.8表示,腔体的几何形状为L/H=6.7,平均腔内温度为230 K.应用商用CFD软件对这个非稳定二维流体进行了仿真.结果表明Strehl比随着波长的增加而增大,但是模糊角却变得相对复杂.证明了波长效应并给出了优化的波长模型λ*=2πσt.     

Probe Method for Measuring the Mach Number and the Flow Vector in Edge Plasma

A planar Mach probe to measure the Mach number and the flow vector with respect to the magnetic field in the edge plasma of e.g. a tokamak is being proposed. The probe is constructed as a double probe and the saturation ion current is measured as function of the angle of rotation of the probe surface normal with respect to the magnetic field. The saturation current as function of the angle is calculated by taking a drifted Maxwellian distribution for ions and an effective size of the probe depending on the angle. From the ratio of the saturation currents for plus and minus biases respectively and from the envelope of the saturation current with respect to the rotation angle, the Mach number and the flow vector can be determined.     

A negative biasing experiment to test numerical calculations basedon a neoclassical L-H transition model in stellarators

Numerical calculations based on Shaing's L-H transition theory in stellarators and performed by Dahi et al. determine viscosity as a function of ion flow speed in the interchangeable module stellarator (IMS). The calculations predict local maximums in viscosity at flows corresponding to Mach numbers -2 and -10, but not elsewhere. The local peaks manifest themselves as jumps in flow speed, and as regions of high radial electric field (E_r) shear. By inducting flows swept between Mach numbers ±5, an electron injection biasing probe revealed a jump at -2, but none at +2, in agreement with the numerical results. A series of flow profiles at constant bias confirmed this result. Altogether, these data agree well with the numerical calculations and provide support for Shaing's L-H transition model as applied to stellarators     

预注冷质的空气涡轮火箭性能分析研究

提出了压气机进口预注冷质的空气涡轮火箭(MIPCC ATR)循环,拓展了ATR在高飞行马赫数下的工作能力。给出了基于工质组分的变比热ATR循环热力学建模方法和工质物性改变时对部件特性的修正方法,基于所推导的准无量纲关系式讨论了循环参数之间的相互作用及其对性能的影响。研究结果表明预注冷质不仅能有效拓宽ATR的工作包线,而且可以提高单位推力,所需付出的代价是比冲会有一定程度的降低。     

喷射器结构改进方法及其CFD分析

给出了一种新的喷射器设计方法:等马赫数梯度法,并对其设计方法进行了详细地介绍。在相同的面积比,相同的工作条件下同传统的一维设计方法设计的喷射器的性能进行了对比。结果表明,与传统设计方法相比,采用本文提出的等马赫数梯度设计的喷射器,可有效消除在扩压室入口处产生的激波,减少由于激波所产生的不可逆损失,使喷射器的性能得到了大幅度地提高。     

Experimental and numerical investigation of the recovery ratio of a wedge-shaped hot-film probe

The recovery ratio of a wedge-shaped hot-film probe was determined in an experimental as well as numerical study, since this information is still unpublished and essential for using the probe in hot-film anemometry. The experiments were conducted at the Khristianovich Institute of Theoretical and Applied Mechanics (ITAM) in Novosibirsk, Russia, and the simulations were performed with StarCCM+, a commercial 2nd order finite volume code. In the analysis, the Mach number was varied between M = 2 and M = 4, and the unit Reynolds number ranged from Re₁ = 3.8?10⁶ to Re₁ = 26.1?10⁶ m^?1, depending on the Mach number. During the experiment, the stagnation temperature was kept constant for each Mach number at a separate value in the range of T ₀ = 289 ± 7 K. Three different stagnation temperatures were used in the simulations: T ₀ = 259 K, T ₀ = 289 K, and T ₀ = 319 K. The difference between the experimental and the numerical results is ≤ 0.5 %, and, therefore, both are in very good accordance. The influence of the Mach number, of the unit Reynolds number, and of the stagnation temperature was analysed, and three different fitting functions for the recovery ratio were established. In general, the recovery ratio shows small variations with all three tested parameters. These dependencies are of the same order of magnitude.     

Rotational temperature and density measurements in a hypersonic flow by dual-line CARS

Rotational temperature and nitrogen number density are measured in the shock wave/boundary layer interaction region in the vicinity of a two-dimensional compression corner disposed in a low pressure, 55 K, Mach 10 hypersonic air flow. Spatially-resolved data are recorded using a nonlinear optical technique named dual-line coherent anti-Stokes Raman scattering (DLCARS). Averaged temperature and density profiles are compared with those predicted by a Navier-Stokes solver. Good agreement is found.