基于POD方法的高超声速边界层转捩判定方法

高超声速边界层转捩是高超声速飞行器设计的关键基础问题之一.为了研究高超声速边界层转捩, 在风洞中, 对平板模型进行了M=5的实验, 在模型中心沿流动方向使用PCB脉动压力传感器对脉动压力时间序列进行采集.文章将本征正交分解(proper orthogonal decomposition,POD)方法引入高超声速脉动压力数据处理中, 发展了单点POD分析方法.经验证, 使用该方法重构数据的均方根(root mean square, RMS)峰值位置可表征转捩位置, 实用性强.      

二维高超声速后台阶表面传热特性实验研究

高超声速后台阶流动是大气层内高速飞行器发动机设计、表面热防护以及高超声速拦截器红外成像窗口气动光学效应校正等诸多先进高超声速技术研发过程中所涉及的一类基础流动问题. 研究高超声速后台阶流动特性对有效提升飞行器综合性能, 进一步掌握高超声速流动机理具有重大基础 意义. 本文以二维高超声速后台阶流动为研究对象, 在KD-01高超声速激波风洞中测量了二维后台阶模型表面传热系数和表面静压, 并将实测台阶下游表面传热系数分布同采用高超声速边界层理论所得估计值进行了比较. 为进一步验证实验结果, 使用NPLS技术测量了其中一种实验状态下台阶周围流动结构. 研究发现, 对于二维高超声速后台阶流动, 台阶下游表面传热分布受台阶处边界层外缘流动特性的直接影响; 在台阶下游分离区和再附区内, 气体黏性占主导作用; 在台阶下游远场区域, 边界层流动特性趋同于平板边界层; 下游边界层基本结构取决于台阶处边界层相对厚度. 对高超声速后台阶流动, 若使用数值模拟方法研究气动热问题, 应当使用湍流模型.     

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...     

高超声速条件下7°直圆锥边界层转捩实验研究

在Ma=6低噪声风洞中开展了半锥角7?的直圆锥边界层转捩相关实验研究.利用响应频率达到MHz量级的高频压力传感器对圆锥壁面脉动压力进行了测量,研究了高超声速圆锥边界层中扰动波的发展过程.结果表明:高超声速圆锥边界层中第二模态扰动波产生的位置以及扰动波特征频率和波长等参数受雷诺数影响较大,当单位雷诺数从2×106m~(-1)增加到8×106m~(-1)时,第二模态波的特征频率从55 k Hz增加到226 k Hz;随着单位雷诺数增加,边界层中扰动增长速度加快,第二模态波出现在圆锥表面更靠近上游的位置;相同单位雷诺数条件下,随着第二模态波的向下游传播,其特征频率逐渐减小.通过对比发现自由来流湍流度对边界层中扰动波的发展同样有较大影响,自由来流湍流度降低,边界层中的第二模态波的特征频率明显减小.利用互相关分析得出第二模态扰动波在边界层中的传播速度大约为当地主流速度的0.8—0.9倍.在1?小攻角条件下,圆锥迎风面和背风面边界层发展呈现出明显的差异,背风面边界层中扰动发展提前,第二模态波出现在更靠近上游的位置,而迎风面中扰动发展受到抑制,第二模态波特征频率更大.     

大后掠钝舵高超声速干扰特性实验研究

文章采用热流率测量和纹影拍摄技术, 对高超声速层流湍流边界层条件下钝舵干扰流场进行了实验研究.实验在高超声速炮风洞内完成, 来流Mach数为6, 8.研究结果表明层流与湍流干扰流场之间存在较大差别, 层流状态下干扰流场存在分离, 平板干扰区内热流率存在负增量, 舵面上存在明显的热流率峰值; 湍流状态下干扰流场无分离, 干扰对平板干扰区内热流率影响较小, 舵面上无明显峰值.      

高速流绕平板边界层特性研究

本项高超声速流绕平板的边界层特性实验研究在中国航天空气动力技术研究院(CAAA)的炮风洞中完成.为了研究分离流动特性, 选择了一项实验研究, 通过实验分别提供绕模型的附着流动与分离流动实验结果.其中第1个模型为顺流平板, 第2个模型为平板上安装突起物, 它们分别对应附着流与分离流动.文章专题研究平板绕流, 为附着流, 它是分离流动的基础.      

三角翼标模摩擦阻力测量的高超声速风洞实验与数据确认

描述了剪切敏感液晶涂层（shear-sensitive liquid crystal coatings,SSLCCs）在高超声速风洞中针对三角翼标模进行表面摩擦阻力（简称摩阻）测量的应用情况.建立了基于剪切敏感液晶涂层光学测量系统,在中国航天空气动力技术研究院FD-07风洞中进行了三角翼摩阻测量实验,给出了三角翼表面摩阻分布数据,表明流场结构的复杂.实验结果证明了此方法进行高超声速摩阻测量的可行性,有效解决了在高超声速流场条件下表面摩阻预测的难题,具有重要的应用价值.      

边界层效应影响等离子体气动激励诱导激波机理

 针对小型暂冲式超声速风洞进行的诱导激波实验结果,提出等离子体气动激励诱导激波的机理不仅取决于放电时产生的焦耳热效应,放电区域的边界层厚度也起到了决定性的作用的推论。为进一步验证此思想,在高超声速激波风洞进行了等离子体气动激励诱导激波的实验研究。结果表明,在边界层很薄的情况下,等离子体气动激励能够诱导出斜激波。分别阐述了两种实验条件下诱导激波的机理,证实了边界层效应在等离子体与激波相互作用中起到了决定性作用。     

高超声速类HTV2模型全目标电磁散射特性实验研究

针对临近空间高超声速飞行器目标探测与识别研究的需求,开展了高超声速飞行器非均匀等离子体电磁散射特性模拟测量研究.利用弹道靶设备发射高超声速类HTV2模型形成模拟的超高速复杂外形目标,弹道靶高精度阴影成像系统和雷达测量系统分别测量高超声速类HTV2模型姿态、全目标C波段/X波段电磁散射特性,获得了不同实验条件下模型全目标雷达散射截面积(RCS)等实验数据.研究结果表明:在不同实验状态下,包覆等离子体鞘套的高超声速类HTV2模型同一测量波段的RCS差别超过1个数量级,模型姿态角对包覆等离子体鞘套的高超声速类HTV2模型RCS影响较大,最大相差1个多数量级;在给定的实验条件下,模型尾迹C波段RCS远小于包覆等离子体鞘套的模型RCS,模型尾迹X波段RCS显著增强;高超声速类HTV2模型全目标C波段电磁散射能量主要分布在模型及其绕流区域, X波段电磁散射能量主要分布在模型及其绕流区域和等离子体尾迹区域.根据弹道靶实验条件,开展了包覆等离子体鞘套的高超声速类HTV2模型电磁散射特性数值仿真,仿真结果与实验结果之间的最大误差小于4 d B,验证了本文提出的非均匀等离子体包覆目标电磁散射特性建模方法的...     

辐射平衡壁温下有限催化模型数值模拟

文章发展了高超声速飞行器辐射平衡壁温下有限催化的数值方法,将数值模拟结果与基于返回舱外形的风洞实验数据进行了对比,并进一步针对典型高超声速飞行器钝双锥研究了辐射平衡壁温下有限催化对气动热环境的影响规律.针对返回舱外形的数值实验表明,完全催化与完全非催化边界条件下壁面热流密度均与风洞实验结果偏差较大,而采用合适的有限催化模型获得的壁面热流密度与风洞实验结果符合良好.针对典型高超声速飞行器钝双锥的研究表明,在辐射平衡温度边界条件下驻点附近氧原子的催化复合系数约为0.17,氮原子的催化复合系数约为0.026,大面积区则分别降为0.005 3和0.01.在驻点热流密度方面,完全催化的壁面热流峰值比有限催化高约21%,而完全非催化的壁面热流峰值比有限催化低约29%.      

超声速平板圆台突起物绕流实验和数值模拟研究

高速飞行器表面不可避免的存在突起物并形成复杂流场, 从而引起飞行器气动特性和热载荷的变化; 同时, 突起物是流动控制的重要方法之一, 合适的突起物形状及安装位置对于改善冲压发动机进气道性能有重要意义. 本文采用基于纳米粒子的平面激光散射技术(NPLS)研究了马赫3.0来流边界层为层流的平板上三个不同高度圆台突起物绕流流场, 主要关注了突起物后方的尾迹边界层, 并采用高精度的显式五阶精度加权紧致非线性格式(WCNS-E-5)离散求解Navier-Stokes方程模拟了该流场. 获得了超声速圆台绕流精细流场结构, 观察到突起物后方尾迹区域边界层发展的过程. 结合实验和数值模拟结果可以发现, 当圆台高度接近或者小于当地边界层厚度时, 突起物对边界层的扰动非常弱, 圆台后方尾迹边界层能够维持较长距离的层流状态, 在边界层转捩阶段也有清晰的发卡涡结构出现; 反之, 边界层受到的扰动明显增大, 在突起物后方很快发展为湍流; 风洞噪声对本文研究圆台引起的边界层扰动有一定影响, 实验获得的边界层转捩位置要比数值结果靠前. 基于NPLS流场图像, 采用间歇性方法分析了圆台突起物后方边界层的特性, 对于高度大于边界层厚度的圆台其间歇性曲线较为接近并且更加饱满, 边界层的脉动也更为强烈.     

Experiments on the unsteady flow field and noise generation in a centrifugal pump impeller

This paper reports on an experimental investigation of large-scale flowfield instabilities in a pump rotor and the process of noise generation by these instabilities. Measurements of the fluctuating components of velocity and surface pressure were made with hot-wire probes and surface mounted pressure transducers on a seven bladed back swept centrifugal water pump impeller operating with air as the working fluid. The impeller was operated without a volute or scroll diffuser, thereby eliminating any sound generation from pressure fluctuations on the volute cutoff. Thus the study focused on flow field and noise components other than the blade passage frequency (and its harmonics). The primary goal of the study was to provide fundamental information on the unsteady flow processes, particularly those associated with the noise generation in the device. It was further anticipated that detailed flow measurements would be useful for the validation of future computational simulations.The measured data at the discharge show a jet-wake type of flow pattern which results in a strong vorticity field. The flow with high velocity found on the pressure side of the impeller tends to move to the low-pressure region present at the suction side of the passage as a form of roll-up around the blade trailing edge. This motion causes an unsteady flow separation at the suction side of the blade and consequently disturbs the flow in the adjacent passage. By interacting with the impeller blades near the trailing edges, this instability flow causes a periodic pressure fluctuation on the blade surface and generates noise by a trailing edge generation mechanism. The spectrum of surface pressure measured at the trailing edge of each blade reveals a cluster of peaks which were identified with azimuthal mode numbers. The correlation between the acoustic farfield pressure and the surface pressure on the impeller blade has proven that the azimuthal modes synchronized with the number of impeller blades generate noise much more efficiently than the other modes. The paper also clarifies the correlation between unsteady flowfield measurements, in both impeller and laboratory co-ordinates, with the radiated noise properties. Thus some light is shed on the noise generation mechanisms of this particular device.     

Investigation of the benefits of unsteady blowing actuation on a 2D wind turbine blade

This paper investigates the benefit of unsteady blowing actuation over a two-dimensional (2D) airfoil specially designed for wind turbine applications. The experiments were carried out in Syracuse University’s anechoic wind tunnel, both with and without large-scale unsteadiness in the free stream generated by a 2D cylinder upstream of the airfoil. By analyzing both surface pressure through wavelet analysis and Particle Image Velocimetry (PIV) velocity field measurements, we found a drastic change in the flow physics and the aerodynamic loading on the airfoil between steady and unsteady free-stream conditions. When there was no large-scale unsteadiness introduced in the flow, under open-loop flow control conditions with unsteady blowing, the leading-edge separation was delayed and the maximum lift coefficient was increased. For the cases where large-scale unsteadiness was introduced into the flow, the experiments showed that both open-loop and closed-loop control cases were capable of reducing load fluctuations by a measurable amount. However, only the closed-loop control case that utilized dynamic surface pressure information from the airfoil suction side near the leading edge was capable of consistently mitigating the fluctuating load.     

基于PXI总线的Φ0.5米高超声速风洞测量系统研制与应用

测量系统是风洞中必不可少的重要的设备之一,其性能对风洞试验的准确性、可靠性和运行效率具有至关重要的影响。近年来,Ф0.5米高超声速风洞原有测量系统逐渐暴露出精准度下降、软件通用性不足、维护不便等缺点,已不能完全满足日益复杂的风洞试验需求。针对此情况,该风洞重新研制建立了一套基于PXI总线的64通道的测量系统,并通过严格测试投入应用,该系统的应用在一定程度上提高了试验数据测量的准确性、可靠性和自动化程度,拓宽了Ф0.5米高超声速风洞的综合性能。     

Pressure-Sensitive Paint measurement of the flow around a simplified car model

In applying Pressure-Sensitive Paint (PSP) to low-speed flow wind tunnel testing, it is important to minimize any measurement uncertainties. There are various error sources such as camera noise, misalignment of images due to model displacement and temperature distribution over the model. Among these factors, the effects of temperature distribution change during tests on pressure measurement accuracies were studied in the present paper. Pressure and temperature distributions over a simplified car model (1/10 scale Ahmed model) were measured using PSP and Temperature-Sensitive Paint (TSP). Sequential images were acquired at the same intervals over the entire test period, including for the conditions before and after the tunnel run. As a result, it was found that the measurement error caused by temperature distribution over the model could be reduced using a single-point temperature measurement. In addition, by measuring surface temperature distributions on the model using TSP, it was proved that the most accurate pressure measurement could be made by rationing the wind-off image acquired immediately after shutting down the tunnel to the wind-on image acquired immediately before shutting down the tunnel. Using the present measurement technique, complicated pressure fields over the Ahmed model were successfully visualized.     

Evolution of disturbances in the shock layer on a flat plate in the flow of a mixture of vibrationally excited gases

The results of the numerical and experimental investigations of the evolution of the disturbances in a hypersonic shock layer on a flat plate streamlined by a flow of the mixture of vibrationally excited gases are presented. The experimental study was conducted in the hot-shot high-enthalpy wind tunnel IT-302 of the ITAM SB RAS. The numerical simulation was carried out with the aid of the ANSYS Fluent package using the solution of the unsteady two-dimensional Navier?Stokes equations with the incorporation of the user-created modules and enabling the consideration of the vibrational non-equilibrium of the carbon dioxide molecules within the framework of the model of the two-temperature aerodynamics. It was obtained that an increase in the carbon dioxide concentration in the mixture with air leads to a reduction of the intensity of pressure disturbances on the surface. The efficiency (up to 20 %) of the method of sound absorbing coatings in the vibrationally excited flows of the mixture of the carbon dioxide and air has been shown.     

钝前缘梯形翼高超声速风洞颤振试验

为了研究钝前缘翼面的高超声速颤振特性,获得典型翼面高超声速颤振参数以校验非定常气动力和CFD计算,采用具有简单结构动力学特性的钝前缘梯形翼模型,在中国航天空气动力技术研究院FD-07高超声速风洞进行了高超声速风洞颤振试验研究.模型为9 mm厚钝前缘梯形平板翼,采用夹层设计:中间层为钢板,提供模型主要刚度和质量特性;两侧为泡沫,起维形作用.试验模型采用悬臂支撑安装于风洞试验段,试验Mach数分别为4.95和5.95.试验固定Mach数,通过缓慢增加动压以使模型达到颤振临界点,采用小波时频谱分析时域响应,结果显示试验模型发生了弯扭耦合经典颤振.试验采用直接观测法获得了颤振动压、颤振频率和对应的试验密度、总温等颤振相关参数.采用壳单元建立了结构有限元模型,并采用统一升力面理论对模型进行了颤振计算分析,研究了气流密度、结构阻尼、Mach数对颤振计算的影响,并对试验结果与理论计算的偏差进行了讨论.分析认为,计算气流密度、计算结构阻尼、结构建模偏差、试验结果散布特性等因素均会构成计算值和试验值之间的偏差,但即便在计算中考虑上述因素,计算结果与试验值仍存在较大偏差.