A long-range laser velocimeter

A long-range laser velocimeter (LV) developed for remote operation from within the flow fields of large wind tunnels is described. Emphasis is placed on improvements in optical hardware as well as additions to data acquisition and processing techniques. The method used for data reduction of photon resolved signals is outlined in detail, and measurement accuracy is discussed. To study the performance of the LV and verify the measurement accuracy, laboratory measurements were made in the flow field of a 10-cm-diameter, 30-m/s axisymmetric jet. The measured velocity and turbulence intensity surveys are compared with measurements made with a hot-wire anemometer. Additionally, the LV was used during the flow calibration of the 80-ft×120-ft wind tunnel to measure the test-section boundary layer thickness at the maximum wind tunnel speed of 51.5 m/s. The requirements and techniques used to seed the flow are discussed, and boundary-layer surveys of mean velocity and turbulence intensity of the streamwise component and the component normal to the surface are presented. The streamwise component of mean velocity is compared with data obtained with a total pressure rake     

Advances in critical technologies for hypersonic and high-enthalpy wind tunnel

Hypersonic and high-enthalpy wind tunnels and their measurement techniques are the cornerstone of the hypersonic flight era that is a dream for human beings to fly faster, higher and further. The great progress has been achieved during the recent years and their critical technologies are still in an urgent need for further development. There are at least four kinds of hypersonic and high-enthalpy wind tunnels that are widely applied over the world and can be classified according to their operation modes. These wind tunnels are named as air-directly-heated hypersonic wind tunnel, light-gas-heated shock tunnel, free-piston-driven shock tunnel and detonation-driven shock tunnel, respectively. The critical technologies for developing the wind tunnels are introduced in this paper, and their merits and weakness are discussed based on wind tunnel performance evaluation. Measurement techniques especially developed for high-enthalpy flows are a part of the hypersonic wind tunnel technology because the flow is a chemically reacting gas motion and its diagnosis needs specially designed instruments. Three kinds of the measurement techniques considered to be of primary importance are introduced here, including the heat flux sensor, the aerodynamic balance, and optical diagnosis techniques. The techniques are developed usually for conventional wind tunnels, but further improved for hypersonic and high-enthalpy tunnels. The hypersonic ground test facilities have provided us with most of valuable experimental data on high-enthalpy flows and will play a more important role in hypersonic research area in the future. Therefore, several prospects for developing hypersonic and high-enthalpy wind tunnels are presented from our point of view.     

激波风洞气动力试验不确定度影响因素分析

参考北大西洋公约组织和AIAA推荐的风洞试验数据不确定度计算方法，结合激波风洞运行特点，确定激波风洞气动力试验的主要误差源，计算激波风洞13-2标模气动力测量结果的不确定度。采用改变单一变量的方法计算主要误差源对测量结果不确定度的影响程度，辨析对不确定度起主要作用的基本参数。计算结果表明：皮托压力和总压的测量结果对流场参数影响显著，皮托压力的测量结果比总压测量结果对流场参数与气动力测量结果影响更大；降低皮托压力和总压的偏离极限，有利于提高激波风洞气动力试验数据的质量。     

油膜干涉测量翼型壁面摩阻低速风洞试验技术

油膜干涉试验可以定量测量得到模型壁面的摩擦应力,是开展摩阻特性研究的有力工具。通过对油膜干涉试验技术中的试验原理、试验方法、误差修正、数据结果分析等关键问题的研究,初步建立了试验系统,具备了开展模型壁面摩阻特性研究的试验能力。综合分析表面热膜试验结果与数值计算结果,在以模型弦长为特征长度,试验雷诺数 Re ＝1．3×106的条件下,翼型表面摩擦应力沿模型弦向的分布规律具有较好的一致性。     

连续式跨声速风洞设计关键技术

为研制先进飞行器,除了提高现有风洞试验测量精度和改进试验技术外,必须建立高性能连续式跨声速风洞试验设备,解决飞行器高速风洞试验模拟能力和精细化模拟问题。以试验段尺寸0．6m×0．6m连续式跨声速风洞设计为例,给出了风洞总体设计方案,分析了如何降低风洞气流脉动、如何改善风洞流场品质、提高风洞运转效率和拓展风洞试验能力等关键技术途径。该风洞作为大型连续式跨声速风洞的引导风洞,方案设计主要采用了高压比压缩机驱动系统、半柔壁喷管、低噪声试验段、高性能换热器和三段调节片加可调中心体式二喉道等新型技术。     

超燃冲压发动机性能参数测量不确定度分析

以AIAA(AIAA S-071A-1999)推荐的不确定度评估标准,借鉴国内外在风洞试验、航空发动机性能测量、火箭发动机地面试验性能等不确定度分析方面开展的研究工作,建立了符合统计学一般原理的超燃冲压发动机总体性能参数测量不确定度分析和计算方法,开展了超燃冲压发动机总体性能参数测量不确定度进行了分析。结果表明:超燃冲压发动机自由射流试验中系统误差引入的不确定度占主导地位,占总不确定度的90%以上。该研究结果有利于提高超燃冲压发动机总体性能测量精度,为超燃冲压发动机的科学研究和工程设计提供支撑。     

基于图像的测量方法在工业风洞跨声速流测量中的最新进展

风洞中非定常复杂流场的实验研究要求先进的测量技术.基于图像的测量技术中最重要的是测量平面流速度场、平面压强分布、模型位置和变形、模型温度以及定量的高速流可视化等技术.DLR(德国宇航研究院)对这些技术的应用包括从低速流到跨声速流、从增升装置到螺旋桨和旋翼、从弹射装置和水塔储水罐尾迹流旋涡到三角翼上涡破裂现象等的研究.由于跨声速风洞的特殊环境,将基于图像的测量技术用于跨声速流要求专门的技术开发和有经验的科学家.给出了DLR空气动力学和流动技术研究所将基于图像的测量技术应用于跨声速流研究的最新进展.     

模型振动对风洞实验数据分散性的数值模拟研究

研究模型振动对风洞实验结果的影响对于修正风洞实验数据和加深空气动力学的理解具有重要的现实意义。基于κ－ω的SST两方程湍流模型,在时间域求解雷诺平均N—S方程,研究了模型振动在不同迎角下流场和气动力的变化规律。结果表明：模型振动在小迎角条件下对气动力测量结果几乎没有影响,但在大迎角存在一定分离的条件下,可以引起分离涡的非线性演化,导致气动力平均值的改变,从而引起风洞气动力测量值的分散性。结论从物理机理上解释了大迎角风洞实验的重复性和数据分散性。     

高超声速风洞气动力试验技术进展

高超声速技术是未来航空航天技术的制高点,而高超声速风洞气动力试验是为高超声速飞行器设计和性能评估提供可靠数据不可或缺的重要技术手段。介绍了高超声速气动力试验设备种类和国内外典型的风洞设备,并分析了目前的发展现状。对国内高超声速风洞气动力试验相关测量技术、试验技术、试验数据评估和高超声速气动力标模体系等研究进展进行了总结。同时,还就高超声速气动力试验设备、气动力试验相关技术的未来发展趋势进行了探讨。     

2.4m跨声速风洞大型飞机试验不确定度评估

2.4m跨声速风洞作为中国目前唯一的大型跨声速气动力试验设备,在中国大型飞机研制中发挥着十分重要的作用。因此,对该风洞试验数据质量的评估、控制和改进提高是一项紧迫的工作。笔者通过完善不确定度计算方法、详细标定基本不确定度源和编制评估软件等工作,建立了该风洞大型飞机试验的不确定度评估方法,并对某大型飞机模型试验结果开展了具体的评估与分析,澄清了该风洞大型飞机试验数据的质量水平。     

大型煤气柜风荷载的风洞试验及数值模拟

对某大型煤气柜进行了风洞测压试验及风压数值模拟。分析了试验模型表面风压分布及其脉动特性,并同数值计算结果、规范条文中类似断面结构的风压分布作对比。结果表明：风洞试验中由于结构表面分布的工字钢及表面粗糙度的处理,雷诺数效应对表面风压分布影响并不明显,但对表面绕流场分离区的风压值有一定影响。结构的均方根升力、阻力系数在频域表现为宽带谱;采用基于雷诺平均的RNGk-ε湍流模型能较准确地模拟表面平均风压分布,其计算结果同样可为结构抗风设计提供参考;在不同量级雷诺数下数值模拟得到的平均风压分布能反映出雷诺数效应的影响。     

多层烧结金属丝网在风洞中的应用

在跨超声速风洞中常采用在稳定段上游安装阻性消声器或谐振腔式消声器来降低暂冲式风洞主回路的主调压阀门、引射器或连续式风洞的压缩机等驱动风洞的动力源所产生的气流噪声强度,达到抑制这些噪声下传影响风洞试验段流场动态品质的效果。随着技术的不断更新特别是降噪技术的不断发展,采用多层烧结金属丝网作为一种新型的消声装置代替原有消声器成为可能。针对这种新技术是否能达到降低风洞气流噪声强度的目的,开展了试验研究方法,通过引导性试验证明,采用多层烧结金属丝网的消声效果优于常规的消声器,且出口气流品质得到较大改善,湍流度大幅度降低,可将这一技术运用到某超声速风洞中。通过性能测试,达到了预期目的,拓展了风洞的降噪技术。     

风洞试验中的视频测量技术现状与应用综述

视频测量（Videogrammetric Measurement,VM）技术因其对试验模型的设计制造无要求,受到国内外风洞试验机构的青睐,本文综述了国外航空航天强国的风洞试验机构所掌握的 VM 技术,并分析了 VM 技术在我国高速暂冲风洞试验中应用所面临的问题,据此中国空气动力研究与发展中心（CARDC）提出了多相机动态标定与基于运动估计的序列图像匹配技术,在暂冲高速风洞高噪声（130 dB 左右）振动环境下,建立了高精度的模型位姿光学测量技术,2米量级高速风洞中的多个应用表明：视频测量的精度高,已用于高速试验模型的姿态与变形测量;另一方面,通过 VM 测量偏折位移场得到光束从摄影中心出发穿过扰流区的光程差,为气动光学效应的研究与测量以及试验模型的壁面流动显示提供新的途径,其光路简单、无需使用价格昂贵的相干光源,因此具有巨大应用前景。     

Aerodynamic characteristics of unsymmetrical aerofoil at various turbulence intensities

A series of wind tunnel tests were performed to investigate the effect of turbulent inflows on the aerodynamic characteristics of the unsymmetrical airfoil at various turbulence intensities and Reynolds number. To assess the aerodynamic characteristics, surface pressure measurements were made over the unsymmetrical airfoil surface by using a simultaneous pressure scanner MPS4264 of Scanivalve make. Self-generated passive grids made of parallel arrays of round bars were placed at four different locations to generate various Turbulence Intensities (TI) in the wind tunnel. The location of the passive grid has been normalized in terms of considering the distance between the entry of the test section and the leading edge of the model. Based on the wind tunnel results, by comparing the baseline without grid low turbulence case TI = 0.51% with other turbulence generated cases like TI = 4.68%, 4.73%, 6.04% and 8.46% at different Reynolds number, it is found that the coefficient of lift increases with the increase in the turbulence intensity. Results also reveal that the flow featuring turbulence can effectively delay the stall characteristics of an airfoil by attaching the flow over the airfoil for an extended region. Additionally, attempts were made to understand the influence of turbulence on the aerodynamic hysteresis.     

大型结冰风洞气流场适航符合性验证

大型结冰风洞气流场适航符合性是大型结冰风洞适航应用的先决条件。为验证3 m×2 m结冰风洞气流场适航符合性,首先建立了结冰风洞气流场适航符合性验证方法,然后针对主试验段构型,开展了气流场适航符合性验证试验,考察了试验段气流速度和喷嘴干空气射流对流场特征参数(气流速度、气流偏角和气流湍流度)的影响,最后评估了试验段内气流场品质,获得了结冰风洞气流场控制包线。结果表明:喷雾耙结构会影响试验段内气流速度和气流偏角空间分布形态,进而导致了非均匀峰值区的形成;增大试验段气流速度会改善气流场品质,但喷嘴干空气射流会显著恶化试验段气流速度低于60 m/s的气流场品质;3 m×2 m结冰风洞主试验段气流场品质在主要试验速度范围内均满足适航审定要求。     

Interplanetary scintillation observations of the high-latitude solar wind

Until the ULYSSES spacecraft reached the polar regions of the solar wind, the only high-latitude measurements available were from indirect techniques. The most productive observations in regions of the solar wind between 5R and 200R have been the family of radio scattering techniques loosely referred to as Interplanetary Scintillation (IPS) (Coles, 1978). Useful observations can be obtained using a variety of radio sources, for example spacecraft beacons, planetary radar echoes and compact cosmic sources (quasars, active galactic nuclei, pulsars, galactic masers, etc.). However for measurement of the high-latitude solar wind cosmic sources provide the widest coverage and this review will be confined to such observations. IPS observations played a very important role in establishing that polar coronal holes (first observed in soft x-ray emission) were sources of fast solar wind streams which occasionally extend down to the equatorial region and are observed by spacecraft. Here I will review the IPS technique and show the variation of both the velocity and the turbulence level with latitude over the last solar cycle. I will also outline recent work and discuss comparisons that we hope to make between IPS and ULYSSES observations.     

电弧加热流场湍流度对尖锥边界层转捩影响的研究

电弧加热流场的热环境特性直接影响热防护系统的地面试验数据,由于电弧加热器高温气流和参数波动的原因,直接测量湍流度非常困难。为研究电弧加热流场湍流度对于边界层转捩的影响,采用红外热图热像仪,在电弧加热流场中进行了5°尖锥模型边界层转捩研究。结合数值计算,将试验结果与常规风洞的尖锥边界层转捩结果进行了比较。结果表明：马赫数影响的雷诺数转捩判别准则可以用于计算电弧加热流场的转捩雷诺数;电弧加热流场的尖锥边界层转捩雷诺数显著小于常规风洞的转捩雷诺数,表明在该试验条件下,电弧加热流场的湍流度显著大于常规风洞。     

常规闭口风洞相阵列气动声学试验

传统气动声学研究观点认为,精确的声学测量要求风洞背景噪声和洞壁反射足够低,传声器测量结果有足够高的信噪比,这是大多数风洞无法达到的要求。近些年,基于声纳和雷达技术发展起来的麦克风相阵列技术可以通过增加阵列的传声器数目从而大幅提高声学测量的信噪比,具有噪声源研究和定位能力,并被成功地应用于非声学固壁风洞噪声源测量和噪声物理机制研究。作者基于相阵列波束生成频域算法研制出常规闭口风洞相阵列系统及相关技术,在FD-09风洞尝试进行了相阵列校准试验和某民机噪声测量试验。结果表明：相阵列技术能够准确捕捉到真实的校准声源,并从技术上验证了相阵列系统在常规闭口风洞测量气动噪声是有效的。     

光学压力敏感涂料测量技术及其在内流场应用

本文介绍了应用基于发光强度的全域压力测量方法进行叶片表面压力分布的一系列实验结果。在自主建立光学压力测量系统和自主研发国产压力敏感涂料的基础上，对高亚音速叶栅风洞出口处大弯度孤立叶片吸力面和对转压气机实验平台出口整流叶片吸力面的压力分布进行了测量，并采用传统电子静压扫描装置在高亚音速叶栅风洞中进行了同步测量。光学压力测量与电子压力扫描结果的对比表明所建立的光学压力测量系统可用于内流场测量，其精度达到了工程应用水平。     

高升力下二维构型阻力准确测量研究

翼型风洞试验阻力测量常使用尾迹流场测量积分求取阻力的方法,但各积分公式均建立在一定的假设基础上,有一定适用范围。在多段翼型流场N-S方程数值模拟和风洞试验的基础上,研究高升力情况下低速风洞阻力精确测量技术。通过N-S方程数值模拟求解多段翼型绕流场,分析尾迹流场的特点和常规风洞试验阻力计算公式推导时所作假设,提出新的更为准确的型阻计算公式;利用多段翼型绕流的数值模拟结果,积分表面压力和摩擦力求得翼型的气动特性,并利用计算得到的尾迹流场信息按照常规和新提出的风洞试验型阻计算公式计算阻力,将三者进行比较,检验提出的新型阻计算公式的准确性;通过风洞试验检验数值模拟得到的流场特点和新型阻计算公式。研究表明：在高升力条件下,传统型阻计算公式有很大的局限性,必须进行改进;提出的考虑尾迹区流动特点的新型阻计算公式能够得到更准确的阻力值。