Airframe noise using experimental, cross‐correlation techniques and a theoretical, vortex model

Airframe noise measurements of a 4.7 percent DC-10 model were recently obtained in the NASA Ames 40 x 80 ft Wind Tunnel. One of the objectives of this test program was to compare model scale measurements with full-scale flyover data to evaluate the scalability af airframe noise from small models. To meet this objective, an array of directional microphones using in-flow parabolic reflectors was designed to increase the signal-to-noise ratio at the scale flyover distance. The on-axis gain of each directional sensor was measured in the wind tunnel during the test using a small speaker source near the model. The directional gain was measured in an anechoic chamber following the test. The measured directivity compared well with predicted values. Measured on-axis gain was lower than predicted, probably due to the losses associated with the short focal length af each reflector necessitated by the in-flow design. Some reverberation effects associated with a covering porous screen were also observed. During the airframe noise test, model generated tones were observed both with near-field sensors and with the far-field reflecting array. Model-generated broadband noise levels were lower than anticipated and below the noise threshold of the directional microphones. The predicted broadband model generated noise at the scale flyover distance is within the measurement range of the multiple microphone phased acoustic array which was successfully demonstrated during the test. (Author)

<div class="htmlview paragraph">There is increasing concern about potential differences in aerodynamic behavior measured in steady flow wind tunnel conditions and that which occurs for vehicles on the road. As tools become available for better simulation of on road conditions there is a growing practical value in understanding what range of conditions are important to simulate.</div> <div class="htmlview paragraph">Surface pressures measured on the sideglass of a European hatchback vehicle in the MIRA full scale wind tunnel are compared with those measured on-road. The on-road data corresponds to relatively calm, low yaw conditions and the time averaged pressure distributions on-road and in the wind tunnel at zero yaw were very similar.</div> <div class="htmlview paragraph">Variations in instantaneous aerodynamic yaw angle produces fluctuations in surface pressures but the sensitivity of instantaneous pressures to yaw angle was lower for the on-road measurements compared with steady state wind tunnel tests. Also, surface pressure unsteadiness was larger than could be attributed to yaw angle fluctuations alone.</div> <div class="htmlview paragraph">The response of instantaneous surface pressures to yaw angle fluctuations is reduced for higher frequency fluctuations as small scale turbulence gusts are less able to modify the flow over the entire vehicle. Examination of the spectral relationship between surface pressure and yaw angle indicates that the impact of yaw angle fluctuations at is reduced by a factor of 2 for reduced frequencies above unity.</div>

Boundary-layer transition measurements have been made on two 5° half-angle cones at M^ « 7 in the Ames 3.5-ft Hypersonic Wind Tunnel and the Langley Variable Density Wind Tunnel. Although there were differences between the measured freestream disturbance scales and pressure fluctuation levels, the choice of consistent locations within the transition region, using either thin-film fluctuation or surface heat-transfer data results in excellent agreement between the transition Reynolds number in both facilities. However, there is an apparent connection between changes in freestream pressure fluctuation levels and movement of boundary-laye r transition location in the two facilities. Nomenclature D — wind-tunnel test section diameter / = frequency p = pressure pc — surface pressure on cone p02 — pitot pressure Re — local Reynolds number based on distance from cone apex Rxx, RZZ — spatial correlation coefficients along the wind-tunnel axis and normal to the axis, respectively ST = Stanton number T0 = stagnation temperature Tw — wall temperature u — velocity p — density 6C = cone half-angle x,z — separation distances along the wind-tunnel axis and normal to the axis, respectively

Numerical simulation of the flow around an air data sensor (ADS), which measures flow angles and Mach numbers using surface pressures on a nose cone, was conducted in this study. Effects of the half-cone angle on the flow angle and Mach number measurements were investigated. Results show that a large half-cone angle achieves high sensitivity of flow angle measurements. Results further demonstrated that a small half-cone angle achieves high-sensitivity of Mach number measurements. To satisfy these conflicting requests, we proposed the use of bi-conic nose cones with two gradients. High sensitivity was achieved for both flow angle measurements and Mach number measurements using this bi-conic nose cone.

The microphone data collected in aeroacoustic wind tunnel test contains not only desired aeroacoustic signal but also background noise generated by the jet or the valve of the wind tunnel, so the desired aeroacoustic characteristics is difficult to be highlighted due to the low Signal-to-Noise Ratio (SNR). Classical cross spectral matrix removal can only reduce the microphone self-noise, but its effect is limited for jet noise. Therefore, an Airflow Background Noise Suppression method based on the Ensemble Empirical Mode Decomposition (ABNSEEMD) is proposed to eliminate the influence of background noise on aeroacoustic field reconstruction. The new method uses EEMD to adaptively separate the background noise in microphone data, which has good practicability for increasing SNR of aeroacoustic signal. A localization experiment was conducted by using two loudspeakers in wind tunnel with 80 m/s velocity. Results show that proposed method can filter out the background noise more effectively and improve the SNR of the loudspeakers signal compared with spectral subtraction and cepstrum methods. Moreover, the aeroacoustic field produced by a NACA EPPLER 862 STRUT airfoil model was also measured and reconstructed. Delay-and-sum beamforming maps of aeroacoustic source were displayed after the background noise was suppressed, which further demonstrates the proposed method’s advantage.

9R7. Aeroacoustic Measurements. - Edited by TJ Mueller (Dept of Aero and Mech Eng, Univ of Notre Dame, 112 Hessert Center, Notre Dame IN 46556-5684). Springer-Verlag, Berlin. 2002. 313 pp. ISBN 3-540-41757-5. $99.00.Reviewed by MG Prasad (Dept of Mech Eng, Stevens Inst of Tech, Hoboken NJ 07030).It is well known that experimental work plays an important role in acoustics. In particular, the acoustical measurements in the presence of airflow is challenging due to test conditions. This book deals with this challenging and important topic in acoustics namely aeroacoustics. The five chapters of the book deal with various aspects aeroacoustic measurements with applications. Chapter 1, on “Microphone measurements in and out of air stream,” deals with the acoustic characteristics of wind tunnels. The discussions include criteria for simulation of aeroacoustic test environment with requirements such as propagation effect, scaling, low background noise levels, etc. Chapter 2 presents “Beam forming in acoustic testing.” The use of beam forming technique and its applications in the modeling of wind tunnel noise with non-acoustic hard walls and free jets are discussed. Several beam forming algorithms with techniques for removing interference is discussed in this chapter. Chapter 3 covers “Aeroacoustic phased array testing in low speed wind tunnels.” A comprehensive treatment of the design and application of acoustic phased array measurement capability for low speed wind tunnels is presented. Various aspects of operation such as calibration, data handling, etc, are discussed. Chapter 4 is on “Source characterization by correlation techniques.” The characterization of acoustical sources using correlation techniques is presented. The chapter discusses various mathematical aspects through cross-correlation and coherence functions. Various examples including correlation measurements between surface pressure and far field acoustic pressure field data are included. The last chapter is “An anechoic facility for basic aeroacoustic research.” The chapter presents the design, construction, and performance evaluation of an anechoic wind tunnel facility for low speed and low turbulence applications. Use of such a facility to experimentally investigate fluid-solid interaction is discussed. A specific wind tunnel facility is described. As an example the use of the facility for a marine propeller response to inflow disturbance is discussed. All chapters are well presented. In summary, Aeroacoustic Measurements is a nice integration of presentations of various recognized contributors. The book deals with important issues in design and performance evaluation of wind tunnel test facilities, measurement techniques, and mathematical aspects. Nice photographs and figures are included. Each chapter has a large number of references. The book is recommended for libraries and all graduate students, researchers, and professionals in acoustics. In particular, those involved in experimental aeroacoustics will find this book very useful.

The present study investigates the differences in the predicted lift for a wind tunnel model from either surface pressures or side wall pressures. The investigated model is the NACA63-418 tested in the Poul la Cour tunnel. The measurements are compared with previous conducted wind tunnel campaigns for a Reynolds number of 3E6. The 3D CFD model of the wind tunnel will be used for investigating details of the flow mechanisms, and to validate assumptions and findings from the experimental data. The CFD simulations show that when the flow is attached the lift based on the side wall pressures results in almost the same lift as calculated from the surface pressures. Around the maximum lift and in the stalled region the flow is not anymore 2D. For the NACA63-418 airfoil the lift based on the side wall pressures are generally highest. The CFD simulations confirm that the method used to predict the lift from the limited side wall pressure strips in the wind tunnel is applicable.

The Andrew Fejer Unsteady Wind Tunnel was modified to add a suction duct on top of the test section to generate a vertical velocity component (cross flow). The suction produced on the top was distributed by the individually controlled louvers at the inlet of the suction duct so that the velocity components in the test section vary both temporally and spatially. Steady and traveling-wave transverse flows were generated in the test section. We present theoretical models for the unsteady flow field generated by this configuration and validate these models with experimental measurements. The results indicated that the transverse gusts generated in the test section were essentially irrotational, and a potential flow model makes relatively accurate predictions of the flow field. The suction-driven approach greatly reduced the turbulence level in the flow field relative to jet-driven cross flow approaches, and demonstrated high levels of repeatability. Traveling waves in 1−cos form were generated and showed evident influence on the flow angle of attack. This paper presents a general method to extract a surface pressure field from a skin friction field in complex flows as an inverse problem, focusing on its application to global luminescent oil-film (GLOF) skin friction measurements. The main technical aspects of this method are described, including the basic equation relating surface pressure to skin friction, variational method, numerical algorithm, and approximate method with the constant boundary enstrophy flux (BEF). The proposed method is evaluated through simulations in the Falkner–Skan flow and the flow over a 70°-delta wing to investigate the effects of the Lagrange multiplier, downsampling rate, noise level, and the value of a constant source term in the approximate method. Further, the approximate method is applied to the skin friction fields obtained by GLOF measurements in the flow over a 65°-delta wing and the square junction flow to obtain the normalized surface pressure fields. The proposed method provides a useful tool to obtain the high-resolution fields of both surface pressure and skin friction by GLOF measurements in complex flows (particularly at low speeds)

<div class="section abstract"><div class="htmlview paragraph">China Automotive Engineering Research Institute Co., Ltd (CAERI) has completed a new vehicle aero-acoustic wind tunnel (AAWT), which is located in Chongqing, China, and has been in operation for 5 years. To help addressing the Chinese vehicle market’s need to improve fuel economy, reduce exhaust emissions, and decrease product development period, the wind tunnel was designed and implemented to achieve a high degree of automation for vehicle testing next to a high aerodynamic and acoustic test accuracy for product development. The CAERI wind tunnel was in operation in June 2019, achieving a top speed of 250 km/h. A 5-belt rolling road system with a long center belt for proper wake simulation is installed inside, a test section with very low static pressure gradient and background noise. Wind tunnel calibration and customized measurement activities can be performed with an overhead traversing system. In the present paper, the main facilities of the AAWT are described next to necessary auxiliary equipment. The aerodynamic and acoustic performances of the CAERI AAWT are summarized. Brief comparisons of wind tunnel flow field parameters to the WLTP certification demand are included. In addition, drag measurement results of one sedan and one SUV are compared with measurements in the FKFS wind tunnel. After employing appropriate wind tunnel correction methods for both wind tunnels, the measurements are in good agreement, demonstrating the accuracy and stability of the CAERI AAWT test results.</div></div>

The flow potentials for axial flow about any arbitrary surface of revolution of easy form are found by a new use of the classical ellipsoidal harmonic solutions of Laplace's Equation in three dimensions. From the flow potentials the surface velocity and pressure are computed in the usual way. The treatment is applicable to the solution of the potential flow about a number of bodies of practical importance. For example, it is equally applicable to airships, to fuselages which are bodies of revolution, to engine nacelles or to submarines which are deeply submerged. In order to compare the theoretical and experimental pressures in a specific case of interest, the surface pressures over the U. S. Airship Akron are computed and compared with the Langley Field measurements on a 1/40 scale model. The agreement is excellent. The longitudinal inertia coefficient is also found, in the case of the application to the Akron, from the flow velocity potential and added to the effective coefficient of the boundary layer as determined from the Langley Field results. The total coefficient so found agrees well with values previously assumed.

Acoustic sensor arrays used in wind tunnels for acoustic measurements are affected by channel and background noise which can be greater than the signal sources. The noise is mainly related to the wind tunnel parameters, interactions between the wind and the wall, and the acoustic sensor. This study combines channel signal filtering with processing of the cross spectral matrix (CSM) diagonal elements to recover the signal of interest from the channel noise with an extended Kalman filter to track the phase difference of the background noise to optimize the coherent background noise. Simulations show that this method effectively reduces the number of side lobes, suppresses the side lobe level, reduces the influence of noise on the beamforming results, and greatly improves the imaging by the acoustic sensor array. 摘要 风洞中使用声传感器进行声学测量, 通道噪声和背景噪声不可避免, 其强度甚至超过信号源。噪声的存在与风洞工作参数、风与壁面的相互作用以及声传感器选择有直接关系。该文根据噪声产生的原因, 通过通道滤波与处理谱交叉矩阵对角线元素相结合来优化通道噪声, 设计扩展Kalman滤波器渐进跟踪背景噪声相位差来优化相干背景噪声。仿真结果表明：所提出的方法可以有效减少旁瓣数量、抑制旁瓣水平, 减少噪声对波束形成结果的影响, 改善声传感器阵列的成像效果。

The modeling requirements for validating Navier-Stokes computations of a high-lift trapezoidal wing are investigated. This wing has a full-span slat and a full-span flap and has been tested extensively in the NASA Langley 14 x 22-ft Wind Tunnel and the NASA Ames 12-ft Pressure Wind Tunnel. Because of the size of the wing, there are significant facility effects in the data from the 12-ft wind tunnel. Computational models of the test facility of differing fidelity are developed and tested. Results are compared with experimental lift, drag, and surface pressures. In the computations a simplified, inviscid model of the test section performs as well as a high-fidelity, viscous test-section model. Computed results generally compare very well with experimental data at all but the highest angles of attack. A comparison of computational results from both free-air and wind-tunnel simulations at the same lift condition indicates that it is necessary to simulate the wind tunnel to perform validation using the 12-ft wind-tunnel experimental data

The Low-Speed Anechoic Wind Tunnel at Sun Yat-Sen University is a premier facility designed for experimental research on aerodynamics and aeroacoustics of aircraft scaled models or components. As a 3/4 closed-circuit wind tunnel with a contraction section nozzle dimension of 0.8 m × 0.6 m, it can be configured either as a closed test section or an open-jet section. The freestream velocity in the closed test section can be adjusted from 5 m/s to 80 m/s, while the openjet section ranges from 10 m/s to 70 m/s. The test section is enclosed within an anechoic chamber with dimensions of 5.5 m × 6 m × 3.5 m to ensure non-reflecting conditions. After the wind tunnel’s construction, comprehensive tests were conducted to evaluate its acoustical and aerodynamic performance. This paper presents the design criteria and characterization of the wind tunnel’s key components, along with the methods and results of benchmarking the flow quality and acoustic performance. This wind tunnel features exceptional flow quality with low turbulence intensity of 0.08% , 1% in the closed test section and open-jet section, respectively. The minimal level of background noise can attain a decibel level of 6.9 dBA under no-flow conditions, and at maximum flow velocity (70 m/s ) in the open-jet test section, the background noise reached 74.9 dB(A).The findings demonstrate that the aerodynamic and aeroacoustic performance indicators of the wind tunnel are comparable to similar facilities worldwide, positioning it as the leading aeroacoustic research facility in the Guangdong-Hong Kong-Macao Greater Bay Area at present.

Surface pressure measurements were taken on a hemisphere-on-cylinder turret in a wind tunnel using pressure sensitive paint and fast response pressure transducers. Four different turret protrusion distances were tested to study the characteristics of the unsteady pressure field on the backside and wake of the turret. Proper orthogonal decomposition was used to identify the dominant spatial surface pressure modes acting on the turret in this parametric study. It was found that the further the turret protruded into the freestream flow, the more the surface pressure field became dominated by spanwise antisymmetric surface pressure distributions resulting from anti-symmetrical vortex shedding at a normalized frequency of approximately StD=0.2. For the case of the partial hemisphere, this anti-symmetrical vortex shedding was essentially absent, insinuating that at some protrusion distance, the surface pressure environment on the turret fundamentally changes. The normalized net force rms was calculated on the turret for each configuration. It was found that the greater the turret protrusion, the greater the net force acting in the spanwise direction.

The continued development of quiet engines is now giving rise to situations where airframe noise is comparable with engine noise at approach. The landing gear is a major contributor to airframe noise, and this paper outlines development and testing of a semi-empirical noise model for predicting the benefit of noise control fairings. The model is based on nondimensional source spectra derived from a data base of full-scale tests on A320 landing gear installed in the DNW wind tunnel and is used to predict the noise reduction potential of a variety of fairings installed on the main and noise gears from an A340 aircraft . The model is found to give good agreement with data and this gives confidence that it may be used as an engineering tool to optimise fairing design. The paper also discusses the way in which the model can provide a framework for incorporating CFD data in the design process and also number of factors which complicate the comparison of noise data from wind tunnel and flyover tests.