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In wind tunnel experiments, the high-precision pressure distribution of models is highly required, but existing measurement methods still have certain shortcomings. In order to obtain the global pressure distribution of the wind tunnel model, this paper assimilated the sparse measured data and numerical calculation data of the wind tunnel experiment by Ensemble Transform Kalman Filter (ETKF), and realized the high-precision reconstruction of the full-space flow field based on the finite measurement points of the model surface. Two-dimensional airfoil RAE 2822 and NACA 0012 were used for experimental verification. Sparse reconstruction of pressure results of RAE 2822 is more consistent with the measured results than the linear theory correction. This effect is especially evident at the shock wave position, and the prediction error of the pressure coefficient is reduced by about 3%. The lift coefficient and moment coefficient of the wing calculated by using the modified ETKF set mean of attack angle and Mach number are less than 1% error from the experimental values. Experiment of NACA 0012 is oriented to the full-field sensing application of wind tunnel experiments and explore the feasibility of pressure reconstruction based on a small number of measurement points. The experimental results show that the relative error of pressure coefficients reconstructed using six measurement points on the wing object surface can be reduced to 2.42%, and the comparison results show that the assimilation effect is highly dependent on the data point locations.

The combustion field is usually a complex system of gas-solid-liquid triphase coupling, and the results obtained from combustion diagnostic can support the researches to improve combustion efficiency and reduce pollutant emissions. In order to make the measurement results more accurate, it is urgent to develop advanced detection methods and detection systems. After several years of development, on-line mass spectrometry has the advantages of high sensitivity, fast analysis speed and wide detection range. It can be used for combustion flow field diagnostics under severe conditions such as high temperature and pressure, and can obtain more comprehensive and sensitive diagnostic information. Firstly, this paper summarized the development of key technologies, e.g., analyzer, ionization source, and sampling systems for on-line mass spectrometry in recent years. Secondly, the applications of on-line mass spectrometry on the measurement of flame component concentration and flame temperature in combustion field were enumerated. On this basis, the challenges and development prospects of the on-line mass spectrometry in complex combustion field diagnostics were summarized, which could provide reference for the relevant researchers.

Against the increasing requirement of large-scale conditions, there is more demand of the measurement instruments. As an efficient approach of measuring cloud droplets, the measurement distance of the rainbow refractometry is generally less than 50 cm owing to the limitation of the aperture of the imaging system. The synthetic aperture rainbow refractometry is proposed to achieve long distance droplet measurement. Laser beams of multiple wavelengths are used to irradiate droplets and multiple rainbow signals are generated. The multiple rainbow signals are collected and combined into one signal and then inversed to obtain the droplet parameters. In this way, the measurement distance of the rainbow refractometry is increased to around 1.5 m, which breaks the aperture limit of the rainbow refractometry and enlarges its applicability. Based on this technique, the multi-wavelength synthetic aperture rainbow refractometer is developed and suitable for parameter measurement of droplets in the central region of large-scale facilities with instruments located outside. The synthetic aperture rainbow refractometry is verified to achieve long distance droplet measurement. The feasibility and accuracy of the multi-wavelength synthetic aperture rainbow refractometer were proved through experimental tests of water and ethanol droplets with different sizes and refractive indices. The sizes of water and ethanol droplets were within the range from 100 to 200 μm. The droplet size and refractive index uncertainties were within 5 μm and 8 × 10−4, respectively. Therefore, the multi-wavelength synthetic aperture rainbow refractometer is no longer limited by the aperture size and can be applied to larger industrial scenarios, achieving the multi-parameters droplet measurement of size, refractive index, and so on.

Molecular Tagging Velocimetry (MTV) and Particle Imaging Velocimetry (PIV) are often used for flow visualization and velocity field imaging. However, the requirement for tracer particles can bring systematic errors to the velocity measurement of PIV method when the tracer particles have poor followability and uneven distribution. In the case of MTV, although particle seeding is not required, the finite fluorescence lifetime of the tracer molecules typically constrains its use only in high-speed and supersonic flows. To develop a velocity field imaging method with no requirements of tracer particles and suitable for low-speed flows, a novel MTV method based on infrared (IR) laser-induced fluorescence is developed and verified in axisymmetric turbulent jet of carbon dioxide. Resonant vibrational transition of the small gas molecule is selectively excited by an infrared pulsed laser to achieve molecular tagging, and the fluorescence distributions of the excited molecules at different instants are then imaged by an infrared camera, from which the velocity distributions are deduced. The effects of the molecular vibrational energy transfer process model, finite fluorescence lifetime, lateral velocity component and molecular diffusion motion on the fluorescence distribution are analyzed to improve the accuracy of the velocity measurement. The proposed method has been successfully verified in carbon dioxide turbulent jets with velocities ranging from 5 m/s to 51 m/s, and the radial distribution of the axial velocity in the main region of the jet is measured. The radial spatial resolution can reach 107 microns, and the velocity distribution is consistent with the theoretical calculation of turbulent jets and previous experimental results. The relative uncertainty of velocity measurement is better than 8%. This method can be used to obtain high-resolution instantaneous velocity imaging of low-speed flow field. Subsequently, by improving the pulse power, excitation efficiency and repetition frequency of the infrared laser, the measurement accuracy, spatial resolution and temporal resolution of this method can be further improved. Therefore, the proposed method bears great potential to provide a quantitative velocity field imaging method in the near-wall flow, micro-scale flow and large gradient flow where it is difficult to introduce tracer particles.

Combustion efficiency is a key pneumatic performance parameter of the aeroengine, and its accurate acquisition is important for improving performance quality of aeroengines, saving fuel, reducing emission and overall performance matching of the combustion chamber. For accurate acquistion of the aeroengine temperature rise combustion efficiency, an inlet reference temperature sensor, an outlet reference temperature sensor and a medium temperature sensor with high accuracy are designed,measurementpoint correction factor is proposed and used based on the combination of the outlet reference sensor and medium temperature sensor, and field calibration of the aeroengine temperature rise combustion efficiency is realized. The results show that the adopted temperature rise combustion efficiency calibration method is reasonable, the calibration result is reliable, and the relative deviation between the temperature rise combustion efficiency measured by the reference temperature sensors and computed by the gas analysis combustion efficiency is 0.3% to 2.1%, and thus the method can solve the calibration problem of the aeroengine temperature rise combustion efficiency.

The interference of the helicopter rotor and the ground causes the lifting of dust, which leads to the phenomenon of brownout. The phenomenon threatens the flight safety of helicopter seriously, and therefore, it is necessary to carry out relevant researches on the aerodynamic basis for the brownout. In this paper, the research progress of helicopter brownout is deeply analyzed from four aspects: numerical methods, experimental techniques, formation mechanism, and suppression methods. Investigations demonstrate that the coupled vortex method or CFD method and the Lagrange sand tracking method can realize simulations of the sand cloud profile and the phenomenon of brownout, but the key factors such as the complex sand bed surface, fluctuating turbulence, migration of sand particles, aggregation, and lifting of sand particles need to be concentrated on. The brownout experimental system built with high-speed PIV techniques and high-speed camera can obtain the morphology data of the sand cloud, but it is still necessary to study the simulation methods of the complex sand bed and the measurement techniques of dust spatial concentration, to explore the driving mechanism of sand particles and the evolution mechanism of brownout, and to develop the design methods and technical approaches to effectively weaken brownout.
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.
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。
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。
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。
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。
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。
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.
采用粒子成像速度场仪(PIV)和数值模拟(CFD)对Taylor-Couette 流场进行测量,获得各转速下涡流场信息。将同等条件下PIV测量结果与数值模拟结果相联系,对比分析不同旋转雷诺数范围内涡流场中不同径线和中轴线上各向速度的变化特征。结果表明,各种特征存在一定的转速分段范围:在2~7r/min(Re为100~350)时,各向速度特征为层流涡特性,在7~40r/min(Re为350~2000)时,各向速度特征为波状涡特性,在40~60r/min (Re为2000~3000)时,各向速度特征为调制波状涡特性,当转速大于60r/min(Re大于3000)时,各向速度特征为湍流涡特性。根据不同角度获得的各向速度特征对应的内筒转速、旋转雷诺数与流场涡形态的关系,明确分析出特定几何条件下,泰勒涡发生形态转变的旋转雷诺数,以便于深入探究泰勒涡流场的特性,定量分析涡运动形态特征。
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。