基于Ludwieg管的高超声速边界层转捩实验

高超声速边界层层/湍流转捩是高超声速飞行器气动力和气动热设计中的难点和热点问题.为了降低开展高超声速边界层不稳定性与转捩实验研究的门槛,研究基于Ludwieg管原理设计并建造了一座Mach 6高超声速管风洞,重点对Ludwieg管风洞的启动和运行过程开展了数值模拟,分析了储气段弯管布局对试验段流场的影响;之后,对该高超声速风洞的自由来流品质进行了静态和动态的标定,验证了风洞的设计Mach数,并给出了流场的动态扰动特征;最后,基于7°半张角尖锥标模开展了高超声速边界层转捩实验,通过表面齐平式安装的高频PCB传感器获得边界层不稳定波,分析了高超声速边界层不稳定波的演化特征.以上工作表明,Ludwieg管相对常规高超声速风洞具有建设和运行成本低、运行效率高、流场品质好等优点,适合开展高超声速边界层转捩等基础实验研究.      

来流Mach数连续变化下的激波串运动特性

为了研究入射激波变化的隔离段内激波串的运动特性,设计并搭建了直连式变Mach数实验系统,捕捉了相同来流Mach数变化速率、不同背压变化速率下激波串的运动行为,揭示了入射激波与背压同时变化对激波串运动的影响机理。入射激波与背压同时变化时,共有3个方面的因素会影响激波串整体的上下游运动趋势,其一,Mach数变化,Mach数增大导致激波串向下游运动,该影响随Mach数增大逐渐减弱;其二,背景波系移动,背景激波反射点靠近激波串前缘时,可能引起激波串的突跳;其三,背压压比变化,背压压比增大时激波串向上游运动,该影响随背压压比增大逐渐增强。三方面因素共同作用下激波串表现出复杂的运动。      

Flow visualization and density measurement of compressible flows by a laser speckle method

Laser speckle method is a well known technique that is useful for both visualization and quantitative measurement. This technique was applied to the density measurement of Mach reflection of shock waves in the present experiment. The Object of the measurement is the density field of simple Mach reflection in relatively low shock Mach number. The non-uniform flow field is divided into three regions by incident, Mach and reflected shock waves. A shock tube was employed in the present experiment. Wedges of 20 degrees and 45 degrees were placed in the test section. YAG laser was employed as a light source. Speckle photograph was taken by a digital still camera. Simple subtraction between the reference and flow images shows a shock pattern and a degree of the correlation of speckle pattern in the flow field. Thus, we can obtain a visualized flow image showing a configuration of Mach reflection from speckle photograph. Speckle photographs which was obtained in the experiments were processed with cross-correlation method. A reconstructed density gradient vector map of Mach reflection was obtained. Comparing the experimental result with numerical one, the measured density gradient shows a good agreement with theoretical prediction.     

Generalized hydrodynamic theory of shock waves: Mach-number dependence of inverse shock width for nitrogen gas

A generalized hydrodynamic theory of shock wave structures is presented for rigid diatomic gases. The theory yields shock solutions for all Mach numbers. The accuracy of the theory is examined for nitrogen gas as an example. Excellent agreement with experiment is obtained for the inverse shock widths for the entire range of Mach number studied by experiment. The comparison made shows that we are now in possession of a successful continuum theory of shock waves in molecular gases for all Mach numbers.     

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.      

空心活塞运动产生的激波

普通活塞在管道中运动, 不论活塞速度高低都能驱动产生激波.空心活塞在管道中运动与普通活塞不同, 当其运动Mach数低于壅塞Mach数, 活塞前不会形成运动激波; 只有活塞运动Mach数超出壅塞Mach数, 才能驱动产生激波.壅塞Mach数由空心活塞中心孔直径、长度和内壁面摩擦系数以及外直径横截面积决定.火车通过隧道类似空心活塞在管道中运动.将火车车箱与隧道间的空隙换算成有效水力直径后, 可利用空心活塞在管道中运动计算壅塞Mach数的方法, 求出火车通过隧道出现激波的条件.文章提出一种新颖的间接测量列车平均摩擦系数新方案.      

强梯度复杂流场中的粒子动力学响应试验研究

示踪粒子在(高)超声速流场中的动力学响应是粒子成像测速等粒子示踪测量技术的关键问题之一.现有文献对粒子动力学响应的试验测量往往是通过单个斜激波响应的测量方法.然而,当示踪粒子用于测量高速飞行器发动机内部复杂的激波串流场时,粒子将经历由多道激波导致的速度、压力、黏性等剧烈变化.本文结合目前(高)超声速飞行器的研究热潮,重点关注示踪粒子在应用于发动机内部具有连续激波的复杂流场测量中存在的跟随性评估方面,开展了一系列的相关试验研究.包括测量超声速风洞的喷管出口速度分布以验证测试系统的性能,在马赫4.2和3.0流场中测量了粒子对二维10°和15°单斜劈绕流中的斜激波动力响应,并测量了模拟发动机内部连续梯度的双斜劈粒子斜激波动力响应.结合粒子动力学的理论模型,得到了各状态的粒子弛豫时间、弛豫距离、Stokes数.基于图像方法、统计学规律分析了激波非定常抖动对测量结果的影响,并对测量结果进行了修正.结果显示,相同斜劈角度下,马赫数越高,粒子的弛豫时间、弛豫距离就越大.但是在相同的来流马赫数下,斜劈角度越大,粒子的弛豫时间、弛豫距离反而减小.在强梯度之后由于流场的雷诺数和黏性系数变化剧烈,粒子的跟随性降低了大约5.7%,stokes数增加了约1%.虽然在本文条件下Stokes数仍满足超声速流场对粒子跟随性的要求,但粒子响应的降低无疑是值得关注的,尤其是当其被应用于具有更多连续梯度的复杂流场测量中.     

高速受限混合层结构和效应实验研究

受限混合层的流动主要是喷流与自由来流相互剪切形成的混合层受到壁面的限制而形成的一种流动.文章采用后向台阶平板模型研究了高速高压比条件下的受限混合层的典型流场结构以及冷却效率.实验自由来流Mach数为5, 喷流的Mach数为1.28, 喷流总压为0.2~0.7 MPa, 通过调整冷喷气流的总压, 基于纹影流动显.形成喷口附近波系的欠膨胀流动现象的深刻认识, 提取波系特征与流动参数之间的规律.基于流动显示及实验测量结果, 通过分析流场中大尺度结构的空间演化规律, 揭示流动参数对于冷却效率的影响规律及物理内涵.采用快响应压敏漆(FRPSP)技术在高超声速风洞开展热流分布和冷却效率研究, 获得了平板对受限混合层冷却效率的影响.      

Flow field study in the T-313 wind-tunnel test section for M = 7

Results of a numerical and experimental study of flow-field characteristics in the test section of the T-313 supersonic blow-down wind tunnel of ITAM SB RAS at Mach number M = 7 are reported. The distributions of local Mach numbers, stagnation temperatures, static pressures, angles of flow deflection from the test-section axis were analyzed. For comparison, distributions of Mach numbers across the flow at several stations at M = 5 and 6 are reported as well. We show that, in the T-313 wind tunnel, two-dimensional nozzle inserts can be used to perform experiments at M = 7.     

高速流场对光传输特性的影响

利用一座小型跨超声速风洞进行了高速流场光传输特性试验研究。光束在高速流场中传输时,由于流场密度变化,光波波前会发生畸变。利用风洞提供0.7,2.0和3.0等气流马赫数的流场条件,采用基于夏克-哈特曼波前传感器的光学测量系统,对光束在风洞流场中传输时的波前畸变进行了测量。试验结果表明:随着风洞流场马赫数增加,流场对光波传播的影响增大,光波波前畸变量显著提高。因此,在利用风洞进行气动光学试验研究之前,有必要消除风洞流场本身对光波传输的严重干扰。     

Infrared spectral radiative properties of high temperature gas with multi-component

The shock layer of high-speed missiles is a highly uneven gas medium with multi-component and high temperature. Its spectral radiation effects degrade and distort the remote sensing images of infrared sensors. The exact numerical calculating method of the shock wave thermal radiation noise has been developed in this paper. With the multi-temperature model and the rectangular grid recursive method of ray tracing, atomic-molecular absorption and emission spectra of CO₂, H₂O and N₂ were obtained along the line of sight along the seeker detectors. Based on image degradation evaluation criteria and the wind tunnel experimental results, how images blurred by the proposed model were verified. The relations between the shock wave thermal radiation noise and flow parameters were also analyzed. For the 3–8 μm infrared band, shock wave thermal radiation noise is little effected by flight altitude, and close relations with Mach number, as the empirical formula given.     

入射激波诱导环氧丙烷爆炸特征研究

用自行设计激波管点火测试技术,实验研究了温度范围760-1380K间入射激波诱导下环氧丙烷的点火机理。利用激波管压力传感器测定了H*(486.1) 和O (470.5nm)随激波诱导强度变化的点火时间特征。实验结果表明：在低马赫数下氢氧自由基出现时间较接近,1.5-2.5马赫间随激波诱导强度增大而线性减小;而马赫大于2.5后,氧自由基的出现时间迅速减小,是由于高活化能的氧自由基的点火时间对强激波较敏感,而诱导强度大于3.5马赫后对两者点火影响区别就下明显了。实验数据将有益于含能材料点火时间的研究。     

入射激波诱导环氧丙烷爆炸特征研究

用自行设计激波管点火测试技术,实验研究了温度范围760-1380K间入射激波诱导下环氧丙烷的点火机理。利用激波管压力传感器测定了H*(486.1) 和O (470.5nm)随激波诱导强度变化的点火时间特征。实验结果表明：在低马赫数下氢氧自由基出现时间较接近,1.5-2.5马赫间随激波诱导强度增大而线性减小;而马赫大于2.5后,氧自由基的出现时间迅速减小,是由于高活化能的氧自由基的点火时间对强激波较敏感,而诱导强度大于3.5马赫后对两者点火影响区别就下明显了。实验数据将有益于含能材料点火时间的研究。     

Dispersive nature of high mach number collisionless plasma shocks: Poynting flux of oblique whistler waves

Whistler wave trains are observed in the foot region of high Mach number quasiperpendicular shocks. The waves are oblique with respect to the ambient magnetic field as well as the shock normal. The Poynting flux of the waves is directed upstream in the shock normal frame starting from the ramp of the shock. This suggests that the waves are an integral part of the shock structure with the dispersive shock as the source of the waves. These observations lead to the conclusion that the shock ramp structure of supercritical high Mach number shocks is formed as a balance of dispersion and nonlinearity.     

激光引致激波演化数值研究

在高重复频率激光推进的研究中,激波的合并是发生在激波演化后期的,同时由于脉冲间隔短,脉冲宽度对流场演化的影响也需要详细研究。考虑了激光辐照过程对流场演化的影响,通过数值计算对激波演化特性进行了研究。结果表明,初期波阵面的椭球形状逐渐转化为一个球形,球心与击穿点的距离随时间逐渐减小并最终趋于稳定。基于激波合并的应用,当激波马赫数在1~2之间时,给出了激波波阵面半径随时间的变化规律,以及激波高压区长度和波峰压强随激波波阵面半径变化规律的经验公式。     

Investigation of the pseudo-shock wave in a two-dimensional supersonic inlet

Abstract  

This paper describes experimental and numerical investigations into the multiple shock waves/turbulent boundary layer interaction in a supersonic inlet. The test model has a rectangular shape with an asymmetric subsonic diffuser of 5°. Experiments were conducted to obtain the visualization images and static pressure data by using supersonic wind tunnel. Numerical simulation was performed by solving the RANS equations with the Menter’s SST turbulent model. The inflow condition was a free-stream Mach number of 2.5 and a unit Reynolds number of 7.6 × 10⁷/m. Numerical results showed good agreement with the experimental results. Based on this agreement, the flow characteristics which are often very difficult to obtain experimentally alone were analyzed with the aid of numerical simulation. The structures, pressure and velocity distributions, and total pressure loss of the pseudo-shock wave in the supersonic inlet were presented in detail from flow visualization images and static pressures.     

Finite-Difference Lattice Boltzmann Scheme for High-Speed Compressible Flow: Two-Dimensional Case

Lattice Boltzmann (LB) modeling of high-speed compressible flows has long been attempted by various authors. One common weakness of most of previous models is the instability problem when the Mach number of the flow is large. In this paper we present a finite-difference LB model, which works for flows with flexible ratios of specific heats and a wide range of Mach number, from $0$ to 30 or higher. Besides the discrete-velocity-model by Watari [Physica A 382 (2007) 502], a modified Lax--Wendroff finite difference scheme and an artificial viscosity are introduced. The combination of the finite-difference scheme and the adding of artificial viscosity must find a balance of numerical stability versus accuracy. The proposed model is validated by recovering results of some well-known benchmark tests: shock tubes and shock reflections. The new model may be used to track shock waves and/or to study the non-equilibrium procedure in the transition between the regular and Mach reflections of shock waves, etc.     

导弹超声速尾退分离干扰特性试验

超/高超声速尾退分离在防热、保形、隐身、多次投放、回收等方面具有明显优势,有望成为高超声速飞行器载荷投放的优选方案。由此面临一类新的多体分离问题:超/高超声速尾退分离问题（aft super/hypersonic ejection separation,ASES）。超/高超声速尾退分离问题本质上是带空腔底部流动与多体分离构成的耦合问题,具有流场结构复杂、气动非定常非线性非对称效应显著的特点。针对超声速尾退分离问题,采用网格测力和轨迹捕获（captive trajectory system,CTS）风洞试验方法探索了尾退分离干扰流场的结构,发现可根据流场结构和舵效变化分为低速-亚声速无激波、高亚声速-跨声速弱激波、超声速激波和准自由流弱干扰4种典型干扰特征,揭示了尾流场影响后不同区域的全弹气动特性和舵效特性以及控制律、攻角、高度和Mach数对分离位移和姿态的影响规律。相关结论将有助于增强对尾退分离问题的认识,对尾退分离技术的工程实践具有参考价值。      

Experimental study of different control methods for hypersonic air inlets

An experimental study of different control methods for hypersonic air inlets aimed at ensuring reliable starting of these apparatuses and improving their operating characteristics in the range of Mach numbers 2 to 8 is reported. Conditions for boundary-layer separation and possibilities for preventing this separation by using modified diffuser configurations and/or perforation bleedage are examined. An air-inlet model was tested for operation in an intermittent wind tunnel and in a blow-down wind tunnel respectively in the Mach-number ranges 2 to 6 and 5 to 8. Distributions of static and total air pressures on the walls of the model and in several cross sections were measured, together with air flow coefficients and total-pressure recovery coefficients. Perforation bleedage is shown to offer an efficient means to facilitate air-inlet starting. Perforation bleed has enabled a more than two-fold increase in the air flow coefficient on the model with sidewalls. A perforation-bleed panel installed closer to the air-inlet throat proved to be more efficient. The possibility of sudden starting of the air-inlet apparatus was checked in the intermittent wind tunnel; it was shown that, here, sudden starting could be realized. The data obtained in the intermittent wind tunnel proved to be consistent with data obtained in the blow-down wind tunnel with up to 150-ms blowdown time. This work was supported by the International Scientific and Engineering Center (Contract No. 887) and by MBDA, France.     

水中斜激波的传播

 利用水的简化物态方程讨论了水中斜激波波后流动马赫数、密度比、压强增量及冲击角与来流马赫数的关系,给出了一个无量纲化的激波极线方程;对附体激波转变为脱体激波及波后流动为等声速流的条件进行了讨论。