考虑对流效应的圆柱绕流气动噪声的预测

采用非稳态雷诺时均(URANS)方程求解圆柱绕流获得近场声源信息,同时运用考虑对流效应的频域声比拟方法,研究了圆柱绕流的气动噪声问题,并考察了不同来流马赫数对辐射声压的影响。结果表明:卡门涡的脱落导致圆柱阻力系数频率为升力系数频率的两倍;较大的升力系数脉动幅值产生了偶极子源,偶极子源在远场辐射声压中占主导作用,最大辐射声压垂直于来流方向;随着马赫数的增大,对流效应显著增强且辐射声压正比于马赫数的三次方。     

多管PDE爆轰噪声传播过程物理特性实验研究

为研究多管脉冲爆轰发动机爆轰噪声传播过程及声波物理特性,对单管至四管脉冲爆轰发动机爆轰噪声物理特性开展实验研究,获得了多管脉冲爆轰发动机爆轰噪声的波形、声压衰减规律、辐射特性、持续时间和频谱特性等物理参数。结果表明:在管口区域,爆轰噪声峰值衰减较快,在远离管口区域,衰减速率逐渐放缓。随着爆轰管数量的增加,爆轰峰值噪声在各方位角上的声压级逐渐增大。在远场区域内爆轰噪声指向性图呈“M”形。随着爆轰管数量的增加,爆轰噪声的A/B持续时间均增加。爆轰噪声是能量主要集中在中低频段的宽频噪声,其频谱峰值和峰值所在频点随爆轰管数量的增加而改变。      

旋涡声散射的空间尺度特性数值研究

旋涡对声波的散射问题是声波在复杂流场中传播的基本问题,在声源定位、声目标识别及探测、远场噪声预测等方面具有重要的学术研究价值和工程应用价值,如飞行器的尾涡识别、探测及测距,湍流剪切流中声目标预测,声学风洞试验中声学测量和声源定位等.声波穿过旋涡时会产生非线性散射现象,其物理机理主要与声波波长和旋涡半径的长度尺度比相关.本文采用高阶精度高分辨率线性紧致格式,通过求解二维非定常Euler方程,数值模拟了平面声波穿过静止等熵涡的物理问题.通过引入声散射截面法,分析了不同声波波长与旋涡半径的长度尺度比对声波脉动压强、声散射有效声压以及声散射能量的影响规律.研究表明:随着声波波长与旋涡半径的长度尺度比逐渐增加,旋涡流场对声场的影响逐渐减弱,声散射有效声压影响区域先逐渐增大随后逐渐减小,声散射能量最大值呈现4种不同的变化阶段.     

高速动车组气动噪声试验与仿真分析*

通过风洞试验对某高速动车组整车、受电弓及转向架远场气动噪声特性进行分析。试验结果表明,高速动车组远场气动噪声是一宽频噪声,总声能随速度的6.6次方增加;由受电弓引起的远场气动噪声主要集中在中高频,噪声峰值频率随速度变化线性增加;由转向架引起的远场气动噪声主要集中在中低频,噪声峰值频率与速度无关。在此基础上,通过大涡模拟和声扰动方程获得该高速动车组近场噪声。高速动车组远场噪声测点仿真结果与试验结果的最大差值2.2 dB(A),最大相对误差2.5%,表明仿真模型的准确性。仿真结果表明,车头近场噪声以车头鼻尖为界,底部气动噪声能量大于上部流线型气动噪声能量,其中转向架舱位置噪声能量最大,因此进行车内外降噪方案设计时,应重点关注车头转向架舱位置。     

转速对弯掠轴流风机气动噪声的影响分析

针对弯掠轴流风机气动噪声问题,采用大涡模拟(LES)和基于Lighthill声类比的FW-H模型相结合的方法进行非定常计算,通过快速Fourier转换(FFT),得到风机远场气动噪声声压级分布。对比研究了三种转速下旋转区内声压级分布规律、时域及频域特性,结果表明:旋转区域内声压级随转速增加而增大,前缘分离涡在某一转速时影响区域和强度最大;在一个旋转周期内,声压脉动呈现出6个波峰与波谷,验证了叶片转动频率是风机内部气动噪声的主要激励频率.     

具有旋涡的翼型绕流辐射声场的数值研究

本文运用面无法和涡声理论数值来研究具有旋涡的ＮＡＳＡＧＡ（Ｗ）－１型翼型绕流的流场和辐射声场，这对于提高透平机械的效率及降低辐射噪声很有帮助。采用作者积累的关于面无法的一些计算技巧，首先成功地求解了不同攻角来流时的流场，再用涡声理论计算了涡旋绕过翼型时声压场的变化及声源的指向性，最后给出了有关这种流动噪声的发声机理和声源特性的一些有用的结论。     

开式轴流风扇气动噪声预测

本文采用LES/FW-H的匹配方法,研究了开式轴流风扇内部旋涡流动特征及其与叶片表面干涉引起的气动噪声之间的联系,同时进行了远场噪声预测,探讨了叶轮不同表面辐射噪声时的频谱分布特征.研究结果表明,开式轴流风扇吸力面附近形成的叶尖涡和前缘分离涡在吸力面叶片表面相应位置形成大压力波动,形成主要噪声源;叶片吸力面的辐射噪声可以通过改善吸力面附近的旋涡流动来降低;低速轴流叶轮由叶轮壁面辐射的噪声以宽频成分为主.     

共轴双旋翼悬停状态气动噪声特性分析

结合CFD(Computational fluid dynamics)方法和FW-H(Ffowcs Williams-Hawkings)方程,建立了一套适合于悬停状态下共轴刚性双旋翼气动噪声特性计算方法。为了准确模拟共轴旋翼流场的涡干扰现象和非定常特性,基于运动嵌套网格技术与双时间推进方法,采用积分形式的可压雷诺平均Navier-Stokes(RANS)方程作为双旋翼非定常流场求解控制方程,湍流模型选用Baldwin-Lomax模型。通过Farassat 1A公式计算双旋翼气动噪声特性,每个声源微面的位置和载荷信息直接从桨叶表面网格中获取。然后,对水平面内和竖直面内观测点处共轴双旋翼厚度噪声、载荷噪声和总噪声的声压时间历程和频谱特性做了细致对比。模拟结果表明:上旋翼和下旋翼反向旋转的特点对声压时间历程影响显著,不同方向观察点的声压波形峰值对应的相位不同;共轴旋翼流场中存在的文丘里效应、桨-涡干扰现象以及下洗流的作用使得桨叶气动载荷呈现明显的非定常特征,导致共轴双旋翼的载荷噪声辐射强度较大;在低频段,总噪声受厚度噪声主导,而在高频段则受载荷噪声主导。      

基于仿生学结构的翼型降噪实验研究

本文运用平面传声器阵列研究了翼型常规尾缘、锯齿形尾缘及波浪形前缘的气动噪声,并在全消声室环境下实验研究了仿生学前缘、尾缘的降噪效果。设计的传声器阵列是同一平面上的单支螺旋线结构,用于二维平面声源定位。实验结果表明:在低湍流度、自由来流情况下,尾缘噪声超过前缘噪声,是翼型噪声的最主要来源,锯齿形尾缘能够有效地降低翼型的尾缘噪声;在较大的宽频范围内,锯齿形尾缘均起到很好的降噪效果,尾缘最大降噪量可以达到10 dB。而波浪形前缘的波长越短,对翼型尾缘的降噪量越明显;波浪形前缘对尾缘噪声的降噪效果在低频时较为明显,而在高频情况下,可以忽略不计。对于湍流-翼型干涉噪声的研究表明:湍流-翼型干涉噪声主要发生于翼型前缘,其声压级远大于尾缘;且锯齿尾缘几乎对湍流-翼型前缘干涉噪声的降噪没有贡献;波浪形前缘可以显著地降低湍流-翼型干涉噪声的声压级,达到良好的降噪效果。     

旋翼桨尖几何形状对旋翼气动噪声影响的定量计算分析

采用CFD与计算声学相结合的Euler Kirchhoff方法对UH-1H直升机旋翼跨声速悬停流动的远场噪声进行定量计算,并将计算结果与实验值进行比较,验证了Euler Kirchhoff方法的正确性.然后以UH-1H直升机旋翼为基准旋翼,定量计算了对桨尖旋翼翼型厚度、旋翼尖削度、后掠角进行变化后的旋翼悬停流场的远场噪声,分析了旋翼桨尖几何形状对噪声的影响.得出结论:降低桨尖翼型厚度、浆尖尖削、浆尖后掠均可以降低噪声峰值.     

Symbolic regression modeling of noise generation at porous airfoils

Based on data sets from previous experimental studies, the tool of symbolic regression is applied to find empirical models that describe the noise generation at porous airfoils. Both the self noise from the interaction of a turbulent boundary layer with the trailing edge of an porous airfoil and the noise generated at the leading edge due to turbulent inflow are considered. Following a dimensional analysis, models are built for trailing edge noise and leading edge noise in terms of four and six dimensionless quantities, respectively. Models of different accuracy and complexity are proposed and discussed. For the trailing edge noise case, a general dependency of the sound power on the fifth power of the flow velocity was found and the frequency spectrum is controlled by the flow resistivity of the porous material. Leading edge noise power is proportional to the square of the turbulence intensity and shows a dependency on the fifth to sixth power of the flow velocity, while the spectrum is governed by the flow resistivity and the integral length scale of the incoming turbulence.     

利用质量流注入的流激孔腔噪声共振峰抑制

通过数值仿真揭示了开口前缘垂直注入质量流和前壁面平行注入质量流抑制流激孔腔噪声的机制,研究了多参数影响下脉动压力峰值降噪量和总降噪量随质量流注入速度的变化规律。开口前缘垂直注入质量流通过抬升剪切层,避免漩涡冲击开口后缘,抑制流激孔腔噪声脉动压力峰值;在一定范围内质量流注入速度越大,脉动压力峰值降噪量越大,但是低频部分引起的抬升也会越高,导致总降噪量先增大后减小;经优化后的峰值降噪量和总降噪量分别可以达到15 dB和9.5 dB。开口前壁面平行注入质量流则是通过加强开口处剪切层的稳定性,避免发生漩涡脱落,达到抑制流激孔腔噪声的目的;当质量流入口面积大于孔腔开口前壁面积2/3时,不仅可以显著降低流激孔腔噪声脉动压力的峰值,并且可以很好地抑制其它频段噪声的抬升;质量流注入速度为来流速度的0.5倍时,脉动压力峰值降噪量和总降噪量分别可以达到18 dB和15.4 dB。     

Some benchmark problems for computational aeroacoustics

This paper presents analytical results for high-speed leading-edge noise which may be useful for benchmark testing of computational aeroacoustics codes. The source of the noise is a convected gust striking the leading edge of a wing or fan blade at arbitrary subsonic Mach number; the streamwise shape of the gust is top-hat, Gaussian, or sinusoidal, and the cross-stream shape is top-hat, Gaussian, or uniform. Detailed results are given for all nine combinations of shapes; six combinations give three-dimensional sound fields, and three give two-dimensional fields. The gust shapes depend on numerical parameters, such as frequency, rise time, and width, which may be varied arbitrarily in relation to aeroacoustic code parameters, such as time-step, grid size, and artificial viscosity. Hence it is possible to determine values of code parameters suitable for accurate calculation of a given acoustic feature, e.g., the impulsive sound field produced by a gust with sharp edges, or a full three-dimensional acoustic directivity pattern, or a complicated multi-lobed directivity. Another possibility is to check how accurately a code can determine the far acoustic field from nearfield data; a parameter here would be the distance from the leading edge at which the data are taken.     

The effect of a cavity on airfoil tones

The presence of a cavity in the pressure surface of an airfoil has been found via experiment to play a role in the production of airfoil tones, which was attributed to the presence of an acoustic feedback loop. The cavity length was sufficiently small that cavity oscillation modes did not occur for most of the investigated chord-based Reynolds number range of 70,000–320,000. The airfoil tonal noise frequencies varied as the position of the cavity was moved along a parallel section at the airfoil's maximum thickness: specifically, for a given velocity, the frequency spacing of the tones was inversely proportional to the geometric distance between the cavity and the trailing edge. The boundary layer instability waves considered responsible for the airfoil tones were only detected downstream of the cavity. This may be the first experimental verification of these aspects of the feedback loop model for airfoil tonal noise.     

An experimental study of airfoil instability tonal noise with trailing edge serrations

This paper presents an experimental study of the effect of trailing edge serrations on airfoil instability noise. Detailed aeroacoustic measurements are presented of the noise radiated by an NACA-0012 airfoil with trailing edge serrations in a low to moderate speed flow under acoustical free field conditions. The existence of a separated boundary layer near the trailing edge of the airfoil at an angle of attack of 4.2 degree has been experimentally identified by a surface mounted hot-film arrays technique. Hot-wire results have shown that the saw-tooth surface can trigger a bypass transition and prevent the boundary layer from becoming separated. Without the separated boundary layer to act as an amplifier for the incoming Tollmien–Schlichting waves, the intensity and spectral characteristic of the radiated tonal noise can be affected depending upon the serration geometry. Particle Imaging Velocimetry (PIV) measurements of the airfoil wakes for a straight and serrated trailing edge are also reported in this paper. These measurements show that localized normal-component velocity fluctuations that are present in a small region of the wake from the laminar airfoil become weakened once serrations are introduced. Owing to the above unique characteristics of the serrated trailing edges, we are able to further investigate the mechanisms of airfoil instability tonal noise with special emphasis on the assessment of the wake and non-wake based aeroacoustic feedback models. It has been shown that the instability tonal noise generated at an angle of attack below approximately one degree could involve several complex mechanisms. On the other hand, the non-wake based aeroacoustic feedback mechanism alone is sufficient to predict all discrete tone frequencies accurately when the airfoil is at a moderate angle of attack.     

A frequency domain numerical method for airfoil broadband self-noise prediction

This paper describes a numerical approach, based in the frequency domain, for predicting the broadband self-noise radiation due to an airfoil situated in a smooth mean flow. Noise is generated by the interaction between the boundary layer turbulence on the airfoil surface and the airfoil trailing edge. Thin airfoil theory is used to deduce the unsteady blade loading. In this paper, the important difference with much of the previous work dealing with trailing edge noise is that the integration of the surface sources for computation of the radiated sound field is evaluated on the actual airfoil surface rather than in the mean-chord plane. The assumption of flat plate geometry in the calculation of radiation is therefore avoided. Moreover, the solution is valid in both near and far fields and reduces to the analytic solution due to Amiet when the airfoil collapses to a flat plate with large span, and the measurement point is taken to the far field.Predictions of the airfoil broadband self-noise radiation presented here are shown to be in reasonable agreement with the predictions obtained using the Brooks approach, which are based on a comprehensive database of experimental data. Also investigated in this paper is the effect on the broadband noise prediction of relaxing the ‘frozen-gust’ assumption, whereby the turbulence at each frequency comprises a continuous spectrum of streamwise wavenumber components. It is shown that making the frozen gust assumption yields an under-prediction of the noise spectrum by approximately 2dB compared with that obtained when this assumption is relaxed, with the largest occurring at high frequencies.This paper concludes with a comparison of the broadband noise directivity for a flat-plat, a NACA 0012 and a NACA 0024 airfoil at non-zero angle of attack. Differences of up to 20 dB are predicted, with the largest difference occurring at a radiation angle of zero degrees relative to the airfoil mean centre line.     

Noise due to turbulent flow past a trailing edge

A theoretical method [I] for calculating far field noise from an airfoil in an incident turbulent flow is extended to apply to the case of noise produced by turbulent flow past a trailing edge, and some minor points of the theory in reference [1] are clarified. For the trailing edge noise, the convecting surface pressure spectrum upstream of the trailing edge is taken to be the appropriate input. The noise is regarded as generated almost totally by the induced surface dipoles near the trailing edge and thus equal, but anticorrelated, noise is radiated into the regions above and below the airfoil wake, respectively. The basic assumption of the analysis, from which these concepts of appropriate input and dominance of dipole sources follow, is that the turbulence remains stationary in the statistical sense as it moves past the trailing edge. The results show that such trailing edge noise often is quite small, compared say to that produced by typical oncoming turbulence levels of one percent, but that it might be appreciable for an airfoil with a flow separation, or for a blown flap.     

柔性旋涡发生器对翼型前缘分离的自适应控制

采用边长为10 mm的三角形柔性和刚性旋涡发生器，安装在二维NACA0018翼型上翼面前缘不同弦长处，用于控制翼型前缘分离流动.实验在低速直流式风洞中进行，以翼型弦长为特征长度的Reynolds数Re=1.1×105，采用单丝热线风速仪测量尾流速度剖面.分别研究柔性和刚性两种材料的三角形旋涡发生器对翼型前缘分离的控制效果.实验结果表明，与刚性旋涡发生器相比，柔性旋涡发生器利用来流能量实现自适应控制，使剪切层下移，从而明显抑制前缘分离.      

Sound radiation from real airfoils in turbulence

Leading edge noise measurements and calculations have been made on a three airfoils immersed in turbulence. The airfoils included variations in chord, thickness and camber and the measurements encompass integral scale to chord ratios from 9 to 40 percent as well as 4:1 ratios of leading edge radius and airfoil thickness to integral scale. Angle of attack is found to have a strong effect on the airfoil response function but for the most part only a small effect on leading edge noise because of the averaging effect of the isotropic turbulence spectrum. Angle of attack effects can therefore be significant in non-isotropic turbulence and dependent on airfoil shape. It is found that thicker airfoils generate significantly less noise at high frequencies but that this effect is not determined solely by the leading edge radius or overall thickness. Camber effects appear likely to be small. Angle of attack effects on the response function of a strongly cambered airfoil are shown to be centered on zero angle of attack, rather than the zero lift angle of attack.