Measuring surface pressure with far field acoustics

This paper presents measurements of the wavenumber frequency spectrum of wall pressure fluctuations under a turbulent boundary layer made using sound radiated from hydrodynamically smooth ridges in the surface. The measurements also serve as a test of the scattering theory of roughness noise. The radiated sound reveals a cut through the full three-dimensional wavenumber frequency spectrum of the wall pressure at the wavenumber of the surface. Since ridges can be made with very small wavelengths, this technique can be used to probe the structure of the wall pressure spectrum on scales far smaller than those that can be reached using conventional wall-mounted transducers. Furthermore, the method reveals the wavenumber frequency spectrum directly, without the need for multi-point measurements or the spatial Fourier transforming of data. Measured spectra bear a close similarity to Corcos’ and Chase's model forms, and confirm the applicability of the theory of roughness noise and its prediction of roughness noise directivity.     

Prediction and test of cavity's hydrodynamic self-noise induced by turbulent boundary layer

The self-noise in cavity is tested in the circling tank,prediction method of cavity's self-noise induced by turbulent boundary layer is established.The window's vibration is using the simply supported boundary condition,the sound wave in the cavity is expanded using the rigid wall boundary condition,the modal coupling vibration equation between them is established using the radiation boundary condition.The turbulent boundary layer pulsating pressure is random,the self-noise power spectrum in the cavity is solved.Test of self-noise and turbulent pressure is carried out in the circling tank when the flow velocity is 5 m/s and 8 m/s,the result verifies that the theoretical method can predict the real cavity's hydrodynamic noise approximately,the trends are similar,this provides one analytical method for sonar dome's material selection and noise control.     

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.     

湍流边界层激励下腔体内水动力自噪声预报与测量

循环水槽中试验测量了腔体内水动力自噪声,并与模态法建立的湍流脉动压力引起腔体自噪声预报进行比较验证。透声窗振动以简支边界为条件,腔体内部声波以刚性边界为条件模态展开,通过辐射边界条件建立模态耦合振动方程。在随机湍流脉动压力作用下,推导了模态振动方程在随机力激励下的自噪声功率谱响应。对循环水槽中5 m/s和8 m/s两种流速工况下的腔体水动力自噪声和湍流脉动压力进行了测试,结合测量的脉动压力预报方法可以计算腔体水动力自噪声量值,理论预报与试验测量结果大致吻合,趋势一致,为声呐罩材料选取及声学环境控制提供了一种分析方法。      

Reduction of turbulent boundary layer induced interior noise through active impedance control

The use of a single actuator tuned to an optimum impedance to control the sound power radiated from a turbulent boundary layer (TBL) excited aircraft panel into the aircraft interior is examined. An approach to calculating the optimum impedance is defined and the limitations on the reduction in radiated power by a single actuator tuned to that impedance are examined. It is shown that there are too many degrees of freedom in the TBL and in the radiation modes of the panel to allow a single actuator to control the radiated power. However, if the panel modes are lightly damped and well separated in frequency, significant reductions are possible. The implementation of a controller that presents a desired impedance to a structure is demonstrated in a laboratory experiment, in which the structure is a mass. The performance of such a controller on an aircraft panel is shown to be effective, if the actuator impedance is similar to but not the same as the desired impedance, provided the panel resonances are well separated in frequency and lightly damped.     

A wave field synthesis approach to reproduction of spatially correlated sound fields

This article discusses an open-loop wave field synthesis (WFS) approach for the reproduction of spatially correlated sound fields. The main application concerns laboratory reproduction of turbulent boundary layer wall pressure on aircraft fuselages and measurement of their sound transmission loss. The problem configuration involves reconstruction of random sound pressure distributions on a planar reproduction surface using a planar array of reproduction monopoles parallel to the reproduction plane. In this paper, the WFS formulation is extended to sound fields with imposed time and spatial correlation properties (or equivalently imposed cross-spectral density in the frequency and wave number domains). Numerical examples are presented for the reproduction of a propagating plane wave, diffuse acoustic field and wall pressure in subsonic or supersonic turbulent boundary layers. The reproduction accuracy is examined in terms of the size of the source plane and reproduction plane, their separation, and the number of reproduction sources required per acoustic wavelength. While the reproduction approach cannot reconstruct sub-wavelength correlation scales of subsonic turbulent boundary layers, it effectively reconstructs correlation scales larger than the acoustic wavelength, making it appropriate for diffuse acoustic field and supersonic turbulent layers.     

On the determination of vertical profiles of temperature, charge concentration, and potential in the surface layer of an aircraft

A “3/2 law” law has been established for the vertical charge concentration profile (i.e., the particle concentration is proportional to temperature to a power of 3/2). The expression is derived for the electric field potential produced by charged particles at the upper boundary of the viscous sublayer, as well as for the density current produced by charged particles moving in the air flow streamlining an aircraft. It is shown that all these parameters increase with altitude and assume maximum values at the turbulent layer surface.     

Experimental estimate of wave spectra of wall pressure fluctuations of the turbulent boundary layer in the subconvective region

A technique is developed for measuring the intensity of the frequency-wave spectrum components of wall pressure fluctuations of the turbulent boundary layer in a quiet aeroacoustic installation with the use of wave filters in the form of rectangular plates. Aluminium-alloy and organic-glass plates of various thickness under a fine-meshed screen are used, set up rigidly flush with the polished wall of the working part of the installation. The experimental data testify to the fundamental possibility of determining the field components of wall pressure fluctuations of the turbulent boundary layer using similar wave filters in the subconvective region, where a substantially lower pressure fluctuation intensity is observed in comparison to the intensity in the region of the convective maximum of the frequency-wave spectrum at a small flow velocity.     

A theoretical model of the pressure field arising from asymmetric intraglottal flows applied to a two-mass model of the vocal folds

A theoretical flow solution is presented for predicting the pressure distribution along the vocal fold walls arising from asymmetric flow that forms during the closing phases of speech. The resultant wall jet was analyzed using boundary layer methods in a non-inertial reference frame attached to the moving wall. A solution for the near-wall velocity profiles on the flow wall was developed based on a Falkner-Skan similarity solution and it was demonstrated that the pressure distribution along the flow wall is imposed by the velocity in the inviscid core of the wall jet. The method was validated with experimental velocity data from 7.5 times life-size vocal fold models, acquired for varying flow rates and glottal divergence angles. The solution for the asymmetric pressures was incorporated into a widely used two-mass model of vocal fold oscillation with a coupled acoustical model of sound propagation. Asymmetric pressure loading was found to facilitate glottal closure, which yielded only slightly higher values of maximum flow declination rate and radiated sound, and a small decrease in the slope of the spectral tilt. While the impact on symmetrically tensioned vocal folds was small, results indicate the effect becomes more significant for asymmetrically tensioned vocal folds.     

Flow-around modes for a rhomboid wing with a stall vortex in the shock layer

The results of theoretical and experimental investigation of an asymmetrical hypersonic flow around a V-shaped wing with the opening angle larger than π on the modes with attached shockwaves on forward edges, when the stall flow is implemented on the leeward wing cantilever behind the kink point of the cross contour. In this case, a vortex of nonviscous nature is formed in which the velocities on the sphere exceeding the speed of sound and resulting in the occurrence of pressure shocks with an intensity sufficient for the separation of the turbulent boundary layer take place in the reverse flow according to the calculations within the framework of the ideal gas. It is experimentally established that a separation boundary layer can exist in the reverse flow, and its structure is subject to the laws inherent to the reverse flow in the separation region of the turbulent boundary layer arising in the supersonic conic flow under the action of a shockwave incident to the boundary layer.     

Estimation of broadband acoustic power radiated from a turbulent boundary layer-driven reinforced finite plate section due to rib and boundary forces

Previous papers considered an infinite fluid-loaded plate with parallel line attachments, driven by a wave-number-white pressure excitation invariant in the direction of an attachment, and established the conditions and procedure for estimating the broadband radiated power by assuming the ribs to radiate independently. This paper applies those results to a finite rectangular ribbed plate, and extends the methodology to include the contribution of the plate's boundary support forces to the radiation and the consideration of excitation that varies in the direction parallel to the ribs. The approach is relevant to problems of sound radiation by underwater stiffened steel plates driven by turbulent boundary layer (TBL) pressures, and is also applicable to stiffened circular cylindrical shells when the response is dominated by bending. Comparisons of sample calculations with results of rigorous models validate the approximation.     

Roles of initial condition and vortex pairing in jet noise

This is a study of the effect of initial condition on sound generated by vortex pairing in a low Mach number, cold air jet (0·15 ⩽ M ⩽ 0·35). Data has been taken, both flow velocity fields and sound pressure far fields, in a quality anechoic facility, with careful documentation of the effect of initial condition on the sound field of jets of two different geometries (i.e., circular and elliptic). Explanations are presented for most of the observed effects by applying Möhring's theory of vortex sound to vortex filament models of coherent structures in the jets. The explanations also draw upon experience with coherent structure dynamics. The sound source of interest here is that associated with the pairing of shear layer vortices. The evolution of these vortices is greatly affected by the initial condition as is their resultant sound field. The elliptic jets with laminar boundary layers show azimuthal directivity, namely, sound pressure levels in the minor axis plane were greater than in the major axis plane. This difference decreases as the nozzle boundary layer undergoes natural transition with increasing jet speed. When the nozzle boundary layer is tripped, making it fully turbulent and removing the shear layer mode of pairing, the elliptic jet sound fields become nearly axisymmetric. What appears to be the most acoustically active phase of vortex pairing has been modeled, and the resulting sound field calculated for the circular jet. Supporting evidence is found in the experimental data for the validity of this model. The model explains the connection between the initial condition and the far field sound of jets. Interestingly, a general result of Möhring's theory is that motions of vortex rings (of any arbitrary shape) can produce only axisymmetric sound fields if the rings remain in a plane. This implies that the observed asymmetric directivity of the laminar elliptic jet sound field must be due to non-planar ring motions of the vortical structures. The primary contribution of this paper is to examine quantitatively the role of vortex pairing in the production of jet noise; the results are used to reemphasize that “pairing noise” cannot be dominant in most practical jet sound fields, contrary to claims by other researchers.     

湍流边界层激励下高速列车车窗参数研究*

高速列车运行过程中,由于结构表面的不平整性产生的湍流边界层激励（TBL）对车厢内部的声场环境产生了一定的影响。采用非相关壁面平面波技术（UWPWs）将随机激励（TBL）转化为一组壁面平面波,利用该激励模型模拟获取了TBL激励下高速列车车窗表面的壁面压力,并将其加载到车窗有限元结构上构建了TBL激励下高速列车车窗模型。分别调查了空腔厚度、双侧玻璃厚度比以及空腔阻尼损耗因子三个参数对车窗声振特性的影响。结果表明：当空腔厚度为20mm,双侧玻璃厚度比为9:5,空腔阻尼损耗因子为0.05时,分别为各自参数下调查工况中的最优解。基于参数调查的结果对车窗结构进行了优化,优化后车窗结构的速度级响应在分析频段内较之前有了显著的削弱,辐射声功率级整体下降了2.8dB。     

Scaling of turbulent wall pressure fluctuations downstream of a rearward facing step

The turbulent boundary layer that forms downstream of a surface discontinuity is locally inhomogeneous. Here, we consider the boundary layer that occurs downstream of a rearward facing step discontinuity on a flat plate that is exposed to a zero incidence, uniform subsonic flow. The wall pressure "point" spectra are measured by small, flush-mounted pressure transducers located at various locations downstream of the step. A new form of the non-dimensional point wall pressure auto-frequency spectrum is proposed that includes the statistical variations of the reattachment location. This form is shown to collapse quite well all of the spectra measured slightly upstream, within, and downstream of the flow reattachment location.     

A synthesis approach for reproducing the response of aircraft panels to a turbulent boundary layer excitation

Random wall-pressure fluctuations due to the turbulent boundary layer (TBL) are a feature of the air flow over an aircraft fuselage under cruise conditions, creating undesirable effects such as cabin noise annoyance. In order to test potential solutions to reduce the TBL-induced noise, a cost-efficient alternative to in-flight or wind-tunnel measurements involves the laboratory simulation of the response of aircraft sidewalls to high-speed subsonic TBL excitation. Previously published work has shown that TBL simulation using a near-field array of loudspeakers is only feasible in the low frequency range due to the rapid decay of the spanwise correlation length with frequency. This paper demonstrates through theoretical criteria how the wavenumber filtering capabilities of the radiating panel reduces the number of sources required, thus dramatically enlarging the frequency range over which the response of the TBL-excited panel is accurately reproduced. Experimental synthesis of the panel response to high-speed TBL excitation is found to be feasible over the hydrodynamic coincidence frequency range using a reduced set of near-field loudspeakers driven by optimal signals. Effective methodologies are proposed for an accurate reproduction of the TBL-induced sound power radiated by the panel into a free-field and when coupled to a cavity.     

Sound pressure in cylindrical shells with regular orthogonal system of stiffeners excited by a random fields of forces

To identify deterministic and random vibrations of a closed cylindrical shell with a regular orthogonal system of stiffeners (stringers and frames) the effective prediction method was worked out. This method is a generalization of the known method of space-harmonic expansion based on the theorem of Bloch-Floquet for a two-dimensional case. The method permits considering correctly the stiffener discreteness and their interaction with the shell through all the components of displacement. On the basis of the method worked out the task of determining the sound field inside the cylindrical volume bounded by the shell excited by the deterministic and random fields of external forces is solved. The high efficiency of the method permits making the prediction of the shell vibrations and of the sound pressure levels inside the volume with account for sound-insulating layers over a wide frequency range. Examples of predicting the strengthened shell vibrations at large aircraft fuselage parameters are given as well as the sound pressure levels inside it under the point force excitation and under excitation by the random field of wall pressure fluctuations of the turbulent boundary layer (within the limits of one of the representation of the space correlation spectrum).     

Wall pressure fluctuation spectra due to boundary-layer transition

Boundary-layer transition has been expected to be an important contributor to sensor flow-induced self-noise. The pressure fluctuations caused by this spatially bounded, and intermittent, phenomenon encompass a very wide range of wavenumbers and temporal frequencies. Here, we analyze the wavevector–frequency spectrum of the wall pressure fluctuations due to subsonic boundary-layer transition as it occurs on a flat plate under zero-pressure gradient conditions. Based on previous measurements of the statistics of the boundary-layer intermittency, it is found that transition induces higher low-streamwise wavenumber wall pressure levels than does a fully developed turbulent boundary layer that might superficially exist at the same location and at the same Reynolds number. The transition zone spanwise wavenumber pressure components are virtually unchanged from the fully developed turbulent boundary-layer case. The results suggest that transition may be more effective than the fully developed turbulent boundary layer in forcing structural excitation at low Mach numbers, and it may have a more intense radiated noise contribution. This may help explain increases in measured sensor self-noise when the sensors are placed near the transition zone. We believe, based on the presented analytical calculation and numerical simulation, that the rapid growth and subsequent decay of turbulent spots in the intermittent transition zone causes the higher low-(streamwise) wavenumber spectra.     

Spatial filtering of wall pressure fluctuations beneath a turbulent boundary layer

Methods of experimental spatial filtering of wall pressure fluctuations beneath a turbulent boundary layer are developed with the aim of obtaining information on the wave number-frequency spectrum. The spatial filtering of the pressure field components by wave-vector filters is studied. The method of spatial filtering of pressure fluctuations by an acoustic array, i.e., a periodic structure with a finite number of elementary transducers, is analyzed. The relation between the wave number characteristic of the acoustic array and the wave number spectrum of the amplitude distribution of transducer’s local sensitivity is determined. Quantitative estimates are obtained for the sensitivity of the array to the wave number spectrum of turbulent boundary-layer pressures, which is necessary for measuring the wall pressure fluctuations in a turbulent boundary layer by wave-vector filters.     

机体表面湍流边界层噪声特性及预测方法研究

湍流边界层噪声是飞机巡航过程中的主要外部噪声源,对舱内噪声水平的影响尤为重要。因此,对飞机机体表面湍流边界层噪声的研究具有重要意义。本文通过试验获得了某型民机巡航过程中的湍流边界层噪声,试飞工况为3500ft/0.78、3500ft/0.7、2500ft/0.67、1500ft/0.66。对实测数据进行分析,发现湍流边界层噪声与动压、边界层厚度等参数有关。同时,利用计算流体力学的方法得到了飞机机体表面的压力分布,并分析了压力梯度对湍流边界层噪声的影响。最后,基于工程预测方法对湍流边界层噪声进行了预测,对于不存在逆压梯度的区域,预测结果与试验结果吻合较好,仅部分频段存在一定偏差。通过对模型的参数进行优化,改善了预测结果。