Despite the decisive influence of various vortex structures of a jet in crossflow on the jet trajectory and scalar mixing, there are few studies related to the high-frequency dynamic characteristics of shear-layer vortexes during transportation. This paper focuses on the high-frequency flow field characteristic, the scalar concentration distribution and the formation and collapse process of the turbulent microstructure of the jet in crossflow with different nozzle diameters and velocity ratios using 40 kHz Particle Image Velocimetry (PIV) and 20 kHz Acetone Planar Laser Induced Fluorescence (Acetone PLIF). The experimental measurements of the velocity and scalar field show that: increasing the velocity ratio promotes the expansion of the circulation zone behind the jet; in the near field of the jet trajectory, power law fitted velocity distribution and shear-layer vortex trajectory shows an exponentially decrease of the jet velocity, the shear-layer vortex strength and vortex motion frequency also show a downward trend, with the frequency of the shear-layer vortex on the windward side slightly lower than that on the leeward side; as the jet velocity increases, the frequency of the shear-layer vortex increases gradually, but the Strouhal number decreases.
The radial spacings of the primary and pilot staged swirler is an important parameter for the lean direct injection combustor. In this paper, the particle image velocimetry technology, Mie scattering technology and particle size measurement technology were used to study the cold flow and spray characteristics under three different radial spacings of the primary and pilot stage. The experimental results show that, under normal temperature and pressure, with the increase of two-stage radial spacing, the central reflux area changes from narrower in front and wider in back to the same width in front and back, the backflow zone between the two stages keeps increasing, the fuel cone angle of the pilot stage is less affected, the main stage jet deflects gradually from the main stage to the pilot stage, and the main fuel crushing effect continues to deteriorate. When the radial spacing of the two-stage is 20 mm, the atomization effect of the main fuel is the best, and the atomization effect of the secondary fuel is also good.
Proper Orthogonal Decomposition (POD) is a reduced order modeling (ROM) method based on 2nd-order statics, which simplifies the investigated wind-pressure field in a new coordinate system formed by a set of orthonormal basis. This paper suggests a method of bi-weighted POD (which weights POD by area and at the same time by root-mean-square), and applies this method to the modal reduction of pressure field around buildings. Firstly, we introduce the POD expansion in a mean-square method, which demonstrates that POD is the optimal choice of ROM in the mean-square sense. Furthermore, we modify the original POD by the bi-weighting-method to improve its capacity of identifying coherent structures with lower energy in pressure field. For the last part, the validity of bi-weighted POD is roughly examined by a case study which applies the method to the pressure field of a 5∶1 rectangular cylinder. It turns out that the modified POD method improves the ROM accuracy at the area associated with lower energy in a significant way. In the meantime, a wind-pressure field ROM constructed by bi-weighted POD captures vital information provided by the original wind-pressure field and is spatially accuracy-consistent.
In cavity structure, complex flows and high-intensity noises appear under the high-speed condition, seriously affecting the aerodynamic characteristics and structural safety of the aircraft. Through the methods of the particle image velocimetry technology and dynamic pressure measurement, the cavity with a length-depth ratio of 3 to 10 is experimentally investigated in the range of Mach number 0.4 to 0.8. The influences of the length-depth ratio and Mach number on the flow field structure in the cavity are emphatically analyzed, and the correlations between the noise intensity and the flow velocity are revealed. The results show that: as the length-depth ratio increases, the thickness of the shear layer in the cavity increases rapidly and expands into the cavity, leading the impact position on the cavity to move down from the back wall to the bottom, and causing the flow type in the cavity to change from open to closed. The increase of the Mach number inhibits the shear layer from expanding into the cavity and induces the main recirculation vortex to move back and the flow type to be open. The amplitude of the overall sound pressure level is positively correlated with the flow velocity in the back of the cavity.
Droplet spreading on a surface is ubiquitous in a variety of applications including aerospace, industry, and agriculture. Majority of these impacts are oblique, while previous studies focused on orthogonal impacts. Oblique impacts cannot be understood directly by previous theories and/or models. Evolution of film formation following a droplet impacting an oblique surface is investigated experimentally. Evolution of the film shape is obtained under various inclination angles and Weber numbers. Based on a new theory of droplet spreading on oblique surfaces, evolution of the film shape is analyzed. It is found that the film shape at small inclination angles can be predicted reasonably, but the error between the predicted maximum lamella width along the inclination direction and the experimental data is relatively big at large inclination angles since the length of the upstream lamella is assumed as a constant in the theory. Modifications of the theory including more detailed analysis of the length of the upstream lamella lead to an analytical model which permits the theoretical determination of the maximum lamella shape. It is shown that the error between the predicted results and the experimental results can be reduced from 61.8% by the previous theory to 3.2%.
In a conventional hypersonic wind-tunnel, pressure fluctuations of the boundary layer on a 7-degree half-angle sharp cone are measured by surface sensors and are analyzed by the linear stability theory. The influences of unit Reynolds numbers and Mach number on the stability and transition position of the boundary layer are studied. The length of the test model is 800 mm and the radius of the head is 0.05 mm. Test unit Reynolds numbers range from 0.49 × 10 7 m–1 to 2.45 × 107 m–1. Test Mach numbers range from 5 to 8. The angle of attack is 0°. The transition position and the energy spectrum distribution of the disturbance wave in the boundary layer are obtained by the quantitative infrared thermography and high frequency surface pressure fluctuation measurement techniques. The frequency and growth rate of the most unstable wave are analyzed by using the linear stability theory. The experimental results show that the fluctuating pressure signal with obvious characteristics of the unstable wave spectrum can be measured in the transition region. The frequency of the pressure fluctuation is close to that of the second mode instability analyzed by the linear stability theory, and the amplitude variation trend is also similar to that of the theoretical analysis. With the increase of the unit Reynolds number, the instability appears earlier, the dominant frequency is increased, and the transition onset moves forward. The unstable wave in the boundary layer contains the first and second modes. When the free-stream Mach number is equal to 5, the transition is caused by the first mode, and when the Mach number is above 6, the transition is attributed to the second mode.
In flight testing, the aeroengine flight thrust is indirectly obtained by the gas generator method. In order to improve the calculation accuracy of the flight thrust, it is necessary to accurately obtain the characteristics of the exhaust system. The laboratory calibration test and numerical simulation research were carried out by using the large bypass ratio separated exhaust system scale model. The results show that: the core nozzle characteristics obtained by the two methods are consistent, and the values are close. When the maximum core nozzle pressure ratio is 1.44, the deviations of the mass flow and the thrust are 0.73% and 0.18%, respectively; the characteristics of the separated exhaust system obtained by the two methods have the same trend and close values. When the max bypass nozzle pressure ratio equals 1.46, the deviations of the mass flow and the thrust are 0.64% and 0.18%, respectively; when the physical model and geometric model of the large bypass ratio separated exhaust system are reasonably simplified, the characteristic deviations of the separated exhaust system obtained by the two methods are in good agreement.
粒子图像测速技术目前已经发展成为实验流体力学领域应用最广泛的非接触激光测试方法之一,为认知复杂流动机理提供直观的流场信息.本文基于超声速流场PIV技术研究实践,针对示踪粒子布撒器设计、粒子松弛特性模型构建、激波流场测试分析、超声速平板湍流边界层结构分析等方面具体问题的研究和认识,从理论、定量化的角度深入分析了应用于超声速流场PIV技术现阶段依然存在的问题.从应用于超声速流场PIV技术的原理出发,针对高速复杂流场的PIV测试现状,总结了应用于超声速流场PIV技术发展过程中的光学部件、示踪粒子及布撒系统所遇到的一系列挑战,以及国内外利用PIV技术在高速复杂流场研究中所取得的成就,针对PIV技术能否适用于高超声速流场的测量做了系统化地探索.并根据实践经验提出了应用于超声速流场PIV技术未来的发展方向:通用的精确的PIV方法不存在,必须从具体研究的流动机理角度改造相应的PIV测试手段.
高速列车进入隧道时,会产生压缩波,压缩波沿隧道内传播至隧道端口后形成向外辐射的微气压波。本文介绍了采用动模型实验平台在200~350km/h速度范围内对60m双向隧道模型的隧道壁面压力波和出口微气压波开展的实验研究。首先分析了实验数据的有效性;其次给出了初始压缩波最大值随时间的衰减变化规律和微气压幅值随实验速度的变化特性;最后研究了流线形头型长度对微气压波幅值的影响。实验结果表明:在实验速度范围内,隧道压力波和出口微气压波无量纲值保持一致,但隧道出口微气压波与流线型头型长度只能定性描述,定量关系难以确定。
在节段模型风洞试验中,两端设置端板可以有效减小端部效应对风压分布的影响,从而保证气流在模型周围的二维流动,其中端板尺寸是影响端板效果的主要参数。为了明确不同尺寸端板对矩形断面气动特性的影响,以桥梁节段模型中最常见的3种宽高比(B/H分别为1、5和10)的二维矩形断面为研究对象,通过刚性模型测压试验,研究了端板尺寸对各模型的气动力、风压分布和斯托罗哈数St的影响。研究结果表明:模型的端部效应不仅仅对端部附近的风压有影响,对中间位置处风压的影响也不容忽视,设置端板是获得准确试验结果的重要保证;随着断面宽高比(B/H)逐渐增大,端部效应影响的程度和范围逐渐减小;随着端板尺寸的增大,模型背风面风压绝对值逐渐增大并趋向一稳定值;抑制端部效应的最小端板尺寸与结构的风迎角有关,风迎角增大,所需的端板也相应增大;有无端板对斯托罗哈数St也有明显影响。
流体推力矢量技术不采用机械偏转,以流动控制方式实现推力转向,有望成为一种更加高效的推力矢量控制方法。目前实现流体推力矢量的主要方法有激波矢量法、双喉道方法、逆流控制方法和同向流方法等,对以上方法选择具有共性的计算与试验数据,对喷管的推力矢量效率、推力损失和流量系数进行了对比分析。结果表明激波矢量方法、双喉道方法和逆流方法能够在大落压比范围内(NPR=1.89~10)实现推力矢量控制,并且具有俯仰/偏航耦合甚至多轴控制的潜力。相比激波矢量法和逆流方法,双喉道和同向流方法在减少推力损失和提高矢量效率上占有优势,不足之处是双喉道方法对喉道进行控制限制了流量系数,而同向流方法的适用落压比范围受到严重限制。为寻求更加高效的矢量喷管技术,国内外相继发展了多种新概念流体推力矢量方法,对每种方法的控制原理、潜在优势和存在的问题挑战进行了探讨,新方法着眼于从喷流出口下游进行控制,对主流的干扰很小,值得深入研究,同时也为流体推力矢量的下一步研究方向提供了借鉴参考。
通过刚性模型测压风洞试验,研究了圆柱的气动阻力、气动升力系数和风压系数随雷诺数的变化规律,从流场分布的角度分析了气动力变化的原因,并研究了雷诺数影响下的流场在圆柱轴向的相关性。结果表明:在亚临界雷诺数区域,在时间平均上流场沿模型两侧呈对称分布,雷诺数对平均阻力系数和流场影响较小,平均升力系数基本为零。在临界雷诺数区域,随着特定区域大负压区的出现,流场不再对称,出现不容忽视的平均升力和脉动升力。在超临界雷诺数区域,随着对称侧大负压区的出现,流场恢复对称状态,平均升力基本消失。雷诺数对流场的轴向相关性有显著的影响。在雷诺数较低时(亚临界区域),卡门涡在轴向上的尺度相对较大,而随着雷诺数的提高,该尺度逐渐减小,各断面流场的相关性降低。
地面风洞试验和飞行试验是研究高超声速飞行器气动加热的主要手段。针对临近空间复杂气动外形高超声速飞行器气动热环境研究的需要,分析探讨了国内气动热试验及测量技术的发展情况。分析了临近空间高超声速飞行器外形特征以及飞行剖面、边界层转捩和气动热环境特性等,进而分析了气动热环境风洞试验模拟理论,介绍了适用于气动热研究的风洞试验设备及其模拟能力,重点讨论了适用于不同类型风洞的热流测量技术发展近况、存在的问题和发展趋势;在以长时间、高热流、高壁温为主要特征的高超声速飞行试验中,无法应用风洞环境下的热流测量技术,因而介绍了目前飞行试验中采用的气动热测量技术,讨论了根据结构温度反辨识表面热流存在的问题,以及热流传感器表面的"冷点效应"、表面催化特性等因素对飞行试验气动热测量的影响,提出了后续工作中应重点研究和解决的临近空间飞行器气动热环境测量技术问题。
火箭冲压组合发动机包含多个工作模态,不同模态灵活组合的优势使其具有宽速域和广空域的工作特点,兼具加速和巡航的优点.火箭冲压组合发动机燃烧室中存在着亚声速、跨声速和超声速共存的流动结构,具有流动速度高、混合时间短、反应强度大、燃烧空间受限和波系结构复杂等特点.围绕火箭射流的强剪切性、燃烧模式的多样性和燃烧过程的动态性,分析了火箭冲压组合发动机的流动与燃烧特征,总结了面向发动机的高速湍流燃烧研究进展,研究了火箭冲压组合发动机中超声速反应混合层的生长特性、燃烧模式与空间释热分布和动态燃烧特性等问题.通过对碳氢燃料详细化学动力学机理的简化、校验,获得了分别适合于工程计算和细致燃烧机理研究的总包反应与框架机理.从火箭射流主导的反应混合层生长模型,宽范围、变来流工作中流动燃烧过程的不确定性和碳氢燃料动力学的简化与加速算法研究出发,提出了火箭冲压组合发动机基础研究中需要突破的问题,为认识发动机中多尺度燃烧机理、优化多模态燃烧组织提供参考.
被动式燃烧诊断技术是利用火焰自发射辐射信息进行燃烧诊断的一项技术,具有非接触、对环境要求不高、系统紧凑、易于实施等特点,在燃烧场在线测量及诊断中具有独特优势。首先,分析了各类燃烧诊断技术的优势及局限;其次,结合华中科技大学煤燃烧国家重点实验室开展的被动式燃烧测量诊断研究工作,从火焰发射光谱、火焰图像处理、热辐射成像技术三个方面介绍了自发辐射燃烧诊断技术的基本原理及研究现状,利用这三种技术,可实现燃烧状态定性分析以及燃烧流场中温度、组分体积分数等燃烧关键信息的定量计算;最后,指出了自发辐射燃烧诊断技术的发展趋势,即:获得更丰富的检测信号、更高的检测分辨率和精度以及更多的检测结果。
采用时间解析PIV(采样频率为1000Hz)在0.55m×0.4m声学风洞中测量了直径D=20mm圆柱后方7.5倍直径、圆柱两侧各3.3倍直径所围成范围内的绕流尾迹在雷诺数Re=2.74×104下的非定常流场。针对PIV获得的速度场数据,进行流场和频谱特性分析,探讨了圆柱绕流尾迹中的平均流场和脉动流场特性,以及旋涡脱落的频率特性。提出了基于速度场之间相关性的相位平均分析方法,系统分析了圆柱上下两侧旋涡交替生成、脱落、发展并耗散的完整演化过程。结果表明:在圆柱后方存在一个低速回流区,其中心0.8D的位置附近是流动结构变化最剧烈的区域;圆柱后方1.9D位置附近是上/下两侧脱落旋涡交汇、耦合的区域,湍流脉动最强;圆柱绕流尾迹中,旋涡脱落频率对应的斯特劳哈尔数稳定在0.2左右;基于速度场之间相关性的相位平均分析方法简单有效,可以准确地识别绕流尾迹中旋涡交替脱落和发展的时空演化过程,在非定常流场测量方面具有普遍推广意义。
高超声速边界层感受性是边界层转捩预测与控制的关键环节,其对高超声速飞行器研究至关重要。目前关于高超声速边界层感受性的实验研究仍然十分匮乏,为了更好地理解高超声速边界层感受性过程并指导该领域的实验研究,文章梳理了近20年来国际上高超声速边界层感受性问题的研究内容,包括对自由流扰动和壁面扰动的感受性,并主要介绍了Fedorov的前缘感受性理论和模态转化机制。最后总结了自由流扰动中感受性的不同发展路径。