Wind Tunnel Wall Research Articles

Generalized, Three-Dimensional Wind Tunnel Wall Correction Using Nonlinear Least-Squares Optimization

A novel three-dimensional technique for correcting wind tunnel measurements for wall interference has been developed. The flow around the test body and its wake is approximated as a series of inviscid singularities, while the wind tunnel walls and model support are approximated using a panel method. A nonlinear least-squares approach with trust-region reflective optimization is then used to obtain the strength and locations of singularities together with the domain boundary panel strengths by fitting to measured wall pressures. The technique is demonstrated using computational fluid dynamics for the test case of a well-documented slender delta wing body with high blockage and validated against both experiments and legacy data. The nonlinear least-squares wall correction method is shown to demonstrate a significant improvement in wall corrections compared to the classical two-variable technique.

Wind tunnel wall interference correction for transonic airfoils with data-reduced ensemble Kalman filter

The uncertain factors in wind tunnel experiments, especially the interference effects, will cause the flow around the test model to differ from that under the free flow condition. The wind tunnel wall interference effects are challenging to be eliminated, especially in transonic flow regions where classical linear correction methods are ineffective. To address this issue, an efficient wind tunnel wall interference correction framework based on data assimilation is proposed for steady and shock buffet flows around the airfoils in the transonic region in this work. The ensemble Kalman filter (EnKF) is used to find the combination of inflow conditions, which minimizes the differences between the pressure coefficients measured in the wind tunnel experiment and those estimated by computational fluid dynamics (CFD). The polynomial chaos expansion method is used to propagate uncertainty in the forecast step of the EnKF in order to improve efficiency. Typical wind tunnel wall interference correction problems are studied, including two steady cases: the transonic flow around the Royal Aircraft Establishment 2822 (RAE2822) and the National Aerospace Laboratory 7301 (NLR7301) airfoils, and one unsteady case: the transonic shock buffet on the National Aeronautics and Space Administration supercritical (phase 2)-0714 (NASA SC(2)-0714) airfoil. It is demonstrated that with the corrected Mach number and angle of attack, the estimated CFD results are in better agreement with the measured experimental data than traditional linear methods and the computational efficiency is improved compared to the original EnKF.

Dynamic Wall Shear Stress Measurement Using Event-Based 3D Particle Tracking

We describe the implementation of a 3d Lagrangian particle tracking (LPT) system based on event-based vision (EBV) sensors and demonstrate its application for the near-wall characterization of a turbulent boundary layer (TBL) in air. The viscous sublayer of the TBL is illuminated by a thin light sheet that grazes the surface of a glass window inserted into the flat wall of the wind tunnel wall. The data simultaneously captured by three synchronized EBV-cameras is used to reconstruct the 3d particle tracks within 0.4 mm of the wall on a field of view of 12.0 mm by 7.5 mm. The velocity and position of particles within the viscous sub-layer permit the estimation of the local, unsteady wall shear stress vector under the assumption of linearity between particle velocity and wall shear stress. Thereby, 2d distributions and higher order statistics of the unsteady wall shear stress are obtained. The systematic underestimation of the spanwise shear stress fluctuation can be explained on the basis of DNS data; a methodology for its mitigation is proposed. The employed EBV hardware coupled with suited LPT algorithms provide data quality on par with currently used, considerably more expensive, high-speed framing cameras.

Transonic buffet simulation using a partially-averaged Navier-Stokes approach

The present article assesses the capability of the partially averaged Navier-Stokes (PANS) method to reproduce accurately the self-sustained shock oscillations, also known as transonic buffet, occurring on airfoils and wings at transonic regime under certain conditions of Mach number and angle of attack. The test case under analysis is an OAT15A unswept wing at Mach number M∞=0.73 and Reynolds number Rec=3×106. The three-dimensional flow is studied by accounting for the wind tunnel walls adopted in the experiments of Jacquin et al. [1] in the simulations. The computations on a large-span, confined configuration reveal a strong three-dimensionality of the flow both before and after the buffet onset. Attention is paid to the comparison with unsteady Reynolds-averaged Navier Stokes (URANS) results, to show the benefits of PANS in resolving flow unsteadiness at different flow resolutions, especially on affordable CFD grids, at limited additional cost. In this context, the role of the mesh metrics and the local turbulence level in the formulation of the model is described, as well as the relation of this latter with the spatiotemporal discretization used for the numerical simulations. The aim is to extend the use of PANS and obtain accurate predictions of flow cases involving shock-wave boundary layer interactions without expensive approaches.

Analysis of different dimensions of a virtual wind tunnel and different wall conditions

Growing interest in improving the performance of vehicles is a subject that allows for continuous research. One such branch of these researchers is energy and consumption performance. These performances describe, on a global level, the need for lower fuel consumption but also a higher energy and power performance. Therefore, to create this balance, direct control of polluting emissions, especially C02 emissions, is necessary. Creating the best mix between low fuel consumption, maximum power of internal combustion engine and low pollutant emissions is influenced by many factors. One such factor, of prime importance, is represented by the aerodynamics of vehicle, in particular the aerodynamic performance, which needs to be tested to be confirmed, so numerical values are needed. Therefore, different ways of testing in the physical or virtual environment are necessary. To facilitate the optimization of a vehicle from an aerodynamic point of view, the most practical way is the virtual wind tunnel, given the possibility to test, at the same time, more configurations. But the CFD (Computational Fluid Dynamics) simulations, sometimes, have the disadvantage of very long compilation times. So, in the presented paper the construction of an optimal virtual wind tunnel for trucks will be presented, focusing on the test area (chamber) and several configurations will be presented such as: different proximity of the tunnel walls to the tested object so a different geometry of the test chamber and condition for the type of floor with different roughness than base case to simulate a more realistic road surface. Following the CFD simulations, it was observed the influence that increasing the dimensions of the test chamber has on the results obtained, so that moving the walls away from the test model generates much more realistic results. In addition, another thing that has an impact is the roughness of the floor, which, in this particular case, generate a 10% difference between the values obtained. Therefore, it will be possible to optimize the tunnel as best as possible to have results as close as possible to reality and in a reduced time.

Relationship between shape of turbulent grid and boundary layer developed on wind tunnel wall in the supersonic wind tunnel

Relationship between shape of turbulent grid and boundary layer developed on wind tunnel wall in the supersonic wind tunnel

Transonic Flutter Dips of the AGARD 445.6 Wing

The AGARD 445.6 configuration is the most popular validation test case for transonic flutter predictions, but the actual extent of truly nonlinear transonic flow for this case is unclear, due to the sparsity of the experimental data and the thin profile of the wing. This work utilizes a combination of mesh adaptation and the linearized frequency-domain method to obtain high-quality viscous and inviscid flutter predictions; these solutions show a double flutter dip through the transonic Mach range driven by complex shock growth across the wing. A single experimental flutter point lies in this flutter dip area, which is not enough to assess the accuracy of these transonic flutter predictions. Modeling the boundary layer of the wind tunnel wall (as opposed to the commonly assumed symmetry wall assumption) has a very large impact on the predicted flutter boundary, though the size of the boundary layer during the original experiment is uncertain.

Evaluation of wind tunnel wall interference using homogeneous and measured boundary conditions

The experimental results obtained in a wind tunnel must be subjected to a correction process which aims to eliminate the influence of the limited dimensions of the flow field around the model. This is necessary because the results must be independent of the characteristics of the laboratory where they were obtained, in order to ensure the quality of the parameters.The wall corrections are applied to the global quantities that characterize the undisturbed flow, such as the Mach number or the dynamic pressure, but they are also applied to the quantities related to the model, namely the global aerodynamic coefficients.Consequently the corrections will be applied to the global quantities of the undisturbed flow and therefore they will be transmitted to the aerodynamic quantities related to the model.

Supersonic Wind Tunnel Flow Characterization Using Planar Laser CO2 Rayleigh Scattering

A complementary experimental and computational study was performed to characterize the test section flowfield of a variable-Mach-number supersonic wind tunnel. A planar laser Rayleigh scattering technique using CO2 nanocrystals was used to capture cross-sectional flow visualizations at various test section streamwise locations. Reynolds-averaged Navier–Stokes simulations of the wind tunnel flow were performed using a k-ω shear stress transport turbulence model to compare against the flow visualizations. Mach 3.20 and 3.55 flow conditions were investigated in the wind tunnel with an empty test section. The Rayleigh signal profile near the wind tunnel walls was compared against boundary-layer thickness data from shadowgraph visualizations, pitot survey measurements, and numerical simulation predictions to guide the interpretation of the scattering results. A brief discussion regarding the nature and potential source of certain unexpected scattering patterns in the images was provided. The most distinguishable flowfield characteristics observed in the Rayleigh scattering visualizations were the asymmetric thickness of the boundary layers along the test section walls and the appearance of two floor lobes near the flow-path corners. The results of the numerical simulations showed favorable agreement with the experimental data and were used to provide insight into the mechanism that caused the boundary-layer features observed. This study demonstrated the usefulness of CO2 Rayleigh scattering to obtain flow visualization data of great value to better understand flow quality in supersonic wind tunnels with two-dimensional nozzles.

Wall Pressure Fluctuations on the Surface of a Fairing in the Shape of a Semi-Ellipsoid and in its Vicinity

The wall pressure fluctuation fields on fairing surfaces and in the surrounding turbulent boundary layer are experimentally investigated. The fairings were in the shape of semi-ellipsoids and were mounted on the wall of a subsonic low-noise wind tunnel. Their heights amounted to 25% of the oncoming boundary layer thickness. The main physical features of the flows under consideration are determined by means of the surface oil-flow visualization. The fluctuating wall pressure field is compared with the flow pattern. It is shown that the greatest pressure fluctuation levels are recorded in the nose region of the fairing surface. It is established that lengthening the model leads to a considerable reduction of the pressure fluctuation strength in the region of flow separation from the fairing surface.

Investigation of a strain-gauge dynamometer for measuring boundary layer friction force on the wind tunnel wall

A one-component strain-gauge dynamometer with a measuring range of 0-0.33 N was developed and manufactured at the Central Aerohydrodynamic Institute to measure directly the unsteady friction force of the boundary layer on the wall of a wind tunnel at very high Reynolds numbers. The subject of this study is to investigate the static and dynamic characteristics of the dynamometer and its systematic errors. The effects of the temperature, longitudinal, normal, and side forces and position of the longitudinal force application were investigated during the calibration. A block was designed and manufactured to calibrate the dynamometer using the dead weights in a wind tunnel. The calibration was carried out in laboratory conditions to determine friction force in the block. The developed dynamometer is a dynamic system with its own natural frequency of oscillations. The correction for the dynamometer's own dynamics is suggested on the basis of the method developed earlier. To calculate the correction to dynamometer dynamics the mass of the metric part of the dynamometer, its natural frequency in the absence of damping and the damping coeffi cient were determined. The correction is verifi ed by experiment with application of a step force. Effect of static and dynamic temperature components on dynamometer readings is investigated. Corrections are proposed to eliminate the systematic errors due to the effect of temperature on the basis of the previously developed method. It is shown that the side force does not affect the dynamometer readings, while the effect of the normal force is 1.1 % of the main sensitivity coeffi cient. The effect of the static temperature component on the sensitivity coeffi cient of the dynamometer is 0.017 %/°C. The total measurement uncertainty of the friction force dynamometer is estimated – the standard deviation of the measurement results does not exceed 0.076 % of the measurement range and the relative standard deviation of the main sensitivity coeffi cient is 0.03 %. The standard deviation of measurement of the nonstationary friction force will not exceed 0.86 % of the range of the dynamometer in the presence of oscillations of the sensitive element at its natural frequency.

Passive Transonic Shock Control on Bump Flow for Wing Buffet Suppression

Since modern transport aircraft cruise at transonic speeds, shock buffet alleviation is one indispensable challenge that civil transport research needs to be addressed. Indeed, in the transonic flow regime shock-induced separation and transonic buffet compromise the flight envelope of an aircraft, and therefore its operational safety and structural integrity. One possible solution is to control and delay the boundary layer separation. The aim of this work was to study whether sub-boundary layer scale period roughness, which locally increases the boundary layer displacement thickness, can act as a virtual shock bump, with aim of bifurcating the foot of the shock wave to reduce the shock’s adverse effect on the boundary layer in the same way as solid shock bumps are known to act. This passive approach can then enhance the buffet margin, consequently extending the safe flight envelope. An experimental investigation was performed, applying this passive technique on a wind tunnel wall bump model which simulated the flow over the upper surface of an aerofoil. The results, in terms of surface pressure distribution and corresponding shadowgraph flow visualisation, showed that such periodic roughness can, indeed, bifurcate the shock wave and delay shock-induced separations, depending on the orientation of the roughness and its periodicity. A virtual shock bump effect can be produced using the displacement effect of periodic sub-boundary layer scale roughness.

Evaluation of the porosity parameter for a perforated wall wind tunnel using measured wall pressure distributions

The existing analytical methods for determining and correcting the effect of wind tunnel walls on experimental data are based on the hypothesis of potential flow. A useful simplification of the boundary conditions used to describe the perforated walls was to consider the wall as homogeneous, the solid and free portions not being treated separately, but as an equivalent permeable surface. The approximation of the wall behaviour during the experiment was possible by defining a porosity parameter. The purpose of this paper is to estimate the porosity parameter for the perforated walls of a trisonic wind tunnel by evaluating the pressure distributions measured on the walls of the test section.

Acoustic wall treatments for wind tunnel aeroacoustic measurements

Sound absorbing porous materials are used to line a wind tunnel wall, in order to reduce reflections. However, the lining can have a detrimental effect on the acoustic measurements due to an increase in the noise radiated from the walls. In addition, the aerodynamic fidelity of the tunnel can be affected. In the present study, the influence of the porous materials on the boundary layer aerodynamic characteristics is assessed. The consequent aerodynamic noise scattering is also studied, and compared against the acoustic benefit from absorbing reflections in the test section. Geometric modelling is used to understand the influence of varying absorbing materials in reducing the acoustic interference caused by the reflections. The aerodynamic and acoustic results are related to the roughness, and to the viscous and inertial resistivities of the three porous materials studied. The material with highest roughness (polyester wool) is found to result in the strongest turbulent fluctuations in the boundary layer. However, it is the material with the thickest fibre diameter (PU foam), and consequent highest inertial resistivity, which generates the strongest surface noise scattering. Materials with high viscous resistivity, together with low inertial resistivity, are found to provide good sound absorbing capabilities. The results therefore indicate that the best choice of sound absorbing wall treatment for wind tunnel applications results from minimizing roughness and inertial resistivity, while maximizing viscous resistivity.

Numerical Modeling of the Walls Perforation Influence on the Accuracy of Wind Tunnel Experiments Using Two-Dimensional Computational Fluid Dynamics Model

The presented work is dedicated to the development of the mathematical model of a closed-type wind tunnel (WT) in a two-dimensional (2D) formulation based on the computational fluid dynamics (CFD) Unsteady Reynolds Averaged Navier Strokes Equations (URANS) method. The influence of the WT walls perforation percentage on the aerodynamic characteristics of the SC1095 (SC is an abbreviation for “supercritical”) airfoil in a wide range of pitch angles from 0 to 28 degrees has been studied. The obtained data are compared with the aerodynamic characteristics of the airfoil in free flow conditions. It is shown that the use of perforation makes it possible to significantly reduce the influence of the wind tunnel walls on the obtained aerodynamic characteristics of the airfoil bringing them closer to the airfoil characteristics under free flow conditions. The developed low resource-intensive 2D CFD WT model can be used to determine the optimal ratio between the walls perforation percentage, the WT test section dimensions, and the experimental model dimensions for various test modes. Support of WT experiments with the help of preliminary parametric studies based on the 2D CFD model could significantly reduce material costs and increase the accuracy of the obtained results.

Effects of windbreak wall model lengths on aerodynamic characteristics of trains on different tracks in wind tunnel tests − a measurement strategy

The length of a windbreak wall model is an important factor in a wind tunnel test. Aerodynamic characteristics of a vehicle on different tracks (windward line and leeward line) with windbreak wall models of different lengths were obtained by wind tunnel tests. The mean pressure and aerodynamic load coefficients (mean drag coefficient, side force coefficient, lift coefficient, and rolling moment coefficient around the lee rail) of the vehicle with the windbreak walls in different lengths were studied. The length of the windbreak wall had a greater effect on the aerodynamic characteristics of the vehicle on the windward line as compared to that on the leeward line. The length of the windbreak wall hardly affected the mean side force coefficient and rolling moment coefficient around the lee rail of the vehicle on the leeward line with the wall in the length over 21.67H (H is the height of the subgrade, including the rail plus the height of the train), whose change was no more than 3.2%. The fluctuations and frequency characteristics of the aerodynamic load coefficients in the length of 21.67H were the same as those with the windbreak wall reaching the wind tunnel wall. A new strategy for measuring the aerodynamic load coefficients of the vehicle with the windbreak wall was presented, which could accurately obtain their means and fluctuations. The maximum error was 1.3%.

Scale effect during the dynamic stall of dynamic airfoil

Studying the scale effect of dynamic airfoil is helpful for understanding the dynamic stall phenomenon and unsteady wall interference. In this paper, the scale effect of the dynamic stall characteristics of the dynamic National Advisory Committee for Aeronautics 0012 airfoil was studied using time-resolved pressure measurements for the first time, which is different from the Reynolds number effect as the experimental Reynolds number was not varied. The evolution of the pressure field and vortex development characteristics on different scaled models was investigated. Significant scale effects were observed on the dynamic airfoil during its dynamic stall. With the model scale increases from c = 500 to 900 mm, the shear layer vortex and the secondary separation were observed, indicating that vortices on the suction surface were enriched. Furthermore, the vortex-induced suction and its occurrence time were significantly affected. The vortex-induced suction varied from Cp ≈ −4.25 to −1.5. With the model scale increases from c = 500 to 900 mm, the vortex structure occurrence angle of attack was delayed by Δα ≈ 0.5°. Further analyses showed that the scale effects of the dynamic airfoil are attributed to the flow transition and unsteady wall interference interactions. Thus, two physical quantities, Sm and Sw, were innovatively defined to associate the vortex complexity, model scale, and wind tunnel wall pressure. This study helps deepen the understanding of the dynamic stall phenomenon and provides guidance for the unsteady wall interference correction.

Acoustic transmission loss and noise from Kevlar wind tunnel walls

In this paper we will develop a model for the acoustic transmission loss and self-noise generated by a Kevlar wind tunnel wall. It is shown that the porosity of the fabric is the most important controlling factor of the transmission loss, and the effect of wind tunnel flow speed is to increase the losses, as observed in experiments. In addition, a model is developed for the weave noise generated by a Kevlar wind tunnel wall, which is found to be caused by the pumping of the fluid through the pores in the Kevlar and depends on their open area ratio. The mechanism for this sound generation is similar to the roughness noise mechanism for a turbulent boundary layer in that the pore spacing couples with the small wavelength disturbances in the boundary layer to cause acoustic radiation at the sum and difference frequencies.

Rotational and blockage effects on a wind turbine model based on local blade forces

This paper describes the results of an extended experimental campaign, reporting surface pressure measurement over one of the blades of the Berlin Research Turbine (BeRT), placed in a closed-loop wind tunnel facility. BeRT is a three-bladed horizontal axis wind turbine with a 3m rotor diameter. The focus is, on the one side, on the three-dimensional effects experienced by the rotating blade, in comparison to 2D approaches by means of XFoil simulations and 2D blade section experiments. On the other side, the blockage effects are investigated between the wind turbine model, placed in the wind tunnel where a 40% blockage ratio is produced, and lifting line free vortex wake simulations, where wind tunnel walls are not considered. Additionally, CFD computations are added in the comparison, with simulations of the far-field and with the wind tunnel walls. The turbine model is studied at several operational conditions such as different blade pitch angles and turbine yaw misalignments. Results are presented in terms of local force components derived from the surface pressure measurements. It is shown that rotational augmentation is evident at the blade mid-span location despite the large blockage. Additionally, the blockage is noticed by means of an offset in both normal and tangential local forces conserving trends and features under axial inflow and yaw misalignments. It is found that the offset in forces can be counteracted by pitching the blades.

Wind Tunnel Installation Effects on Japan Aerospace Exploration Agency’s Standard Model

This paper investigates the interference effects of wind tunnel floor and walls on the Japan Aerospace Exploration Agency’s High-Lift Configuration Standard Model (JSM) as an extension to a paper published in the Journal of Aircraft, Vol. 56, No. 3, May–June 2019, pp. 1080–1098. An unstructured hybrid grid generator, MEGG3D, was applied to JSM wing–body and wing–body–nacelle–pylon configurations to generate grids for free-air flow conditions. These grids were then reused in an investigation of the interference effects of Japan Aerospace Exploration Agency’s low-speed wind tunnel walls to better compare the computational results, both with each other and with experiment. The geometry of the wind tunnel used in our computational simulations is described in detail so that they can be reproduced. A Reynolds-averaged Navier–Stokes flow solver for unstructured hybrid grids, the Tohoku University Aerodynamic Simulation (TAS) code, was used for the simulations. The computational results indicated that the standoff and test section floor boundary layer significantly affected the aerodynamics of the JSM at high angles of attack. Better agreement with experiment was observed for the wing–body configuration in the aerodynamic characteristics near maximum lift if the effects of the standoff and floor boundary layer were considered.