European Transonic Windtunnel Research Articles

A numerical approach for assessing slotted wall interference using the CRM model at ETW

This paper is devoted to the assessment of wall interference in the slotted wall test section of the European Transonic Windtunnel (ETW) over a wide range of Reynolds numbers. The experimental part of the investigation was performed in February 2014 by testing the NASA Common Research Model mounted on a fin-sting support. These tests were carried out within the scope of the ESWIRP project funded by the European Commission in the 7th framework program. The numerical research was based on the Electronic WindTunnel (EWT-TsAGI) software with a cryogenic solver. The assessed Mach number influence on the wall signatures revealed a very similar effect to applying the classical Prandtl–Glauert rule over the investigated Mach number range. Practically, no Reynolds number effects on the wall pressure distributions generated by the model and its support system could be identified over the wide range of Re numbers investigated. The first attempt of the EWT-TsAGI code application for a simulation of ETW tests featuring the model in the slotted wall tunnel showed a fair coincidence of the pressure coefficient distribution on test section walls in the model region, on the wing-root sections and the drag polar at moderate lift coefficient values.

Development of flow structures in the near-field wake region of the Common Research Model

The flow in the near-field wake region of a generic transport aircraft configuration has been studied using both unsteady RANS and delayed detached-eddy simulations. The NASA Common Research Model has been used for this purpose, as experimental data obtained with this model in the cryogenic European Transonic Wind Tunnel during the ESWIRP test campaign were available for comparison and validation. Focusing on high angles of attack of \(\alpha = 16^\circ\) and \(\alpha = 18^\circ\), the results form the basis for the study of flow phenomena occurring at stall conditions. Such flight conditions are characterized by massive flow separation at the wing and highly unsteady flow in the wake. In a first step, the numerical results are compared with available test data using global force coefficients and pressure distributions on the wing. Both show a good agreement of numerical and experimental results, indicating slight deviations of the pressure variable at the inboard wing sections. In a second step, the turbulent structures of the near-field wake are examined using the proper orthogonal decomposition method based on snapshots (two-dimensional instantaneous flow fields). Frequencies of the first mode-pair match the dominant frequencies of the lift and velocity wake spectra very well.

Microphone Array Measurement in European Transonic Wind Tunnel at Flight Reynolds Numbers

This paper presents an experimental investigation of small-scale model airframe noise at various Reynolds numbers up to the real flight case. The study consists of data acquired with a microphone array in the European Transonic Wind Tunnel. The advantage of this wind tunnel is to enable testing simultaneously at cryogenic temperatures and increased pressure levels, which extends the range of achievable Reynolds numbers up to those pertaining to full-scale flight conditions. At the DLR, German Aerospace Center, the microphone array measurement technique has been further developed to perform measurements under combined cryogenic and pressurized conditions. For this purpose, a microphone array consisting of 96 microphones was designed and constructed. In this paper, aeroacoustic results are presented for various Reynolds numbers up to the real flight case on an Airbus K3DY half-model of scale . The results showed significant Reynolds number and Mach number dependency for various sources. Of particular note are various dominant sources appearing on the flap at real flight Reynolds numbers. To the author’s knowledge, this is the first time that airframe noise data for a small-scale model have been acquired at real flight Reynolds numbers.

Model Deformation Corrections of NASA Common Research Model Using Computational Fluid Dynamics

Computational fluid dynamics-aided wind-tunnel data corrections for the static model deformation effect of the NASA-Common Research Model are performed for wind-tunnel data obtained at the Transonic Wind Tunnel of Japan Aerospace Exploration Agency, the National Transonic Facility, and the European Transonic Wind Tunnel. Although it is important to use the same configuration when comparing wind-tunnel data obtained at various facilities, this has not yet been performed with an appropriate validation. In this paper, the model deformation effects are estimated from the difference between the computational fluid dynamics results for a designed configuration and that for a deformed configuration considering support systems. Initially, the techniques for the model deformation correction are validated with two sets of National Transonic Facility wind-tunnel data at a Reynolds number of with different dynamic pressures. Then, the validated model deformation corrections are performed on the wind-tunnel data obtained at the Transonic Wind Tunnel, National Transonic Facility, and European Transonic Wind Tunnel. Differences in the drag coefficients among the three sets of wind-tunnel data fall within about 10 drag counts for the model deformation corrections and the Reynolds number corrections applied to the Transonic Wind Tunnel data to align the Reynolds number to .

Unsteady Wake of the NASA Common Research Model in Low-Speed Stall

Following the completion of the European Strategic Wind Tunnel Improved Research Potential test campaign in the European Transonic Windtunnel, a large amount of data exists for comparison and validation of computational results at high Reynolds numbers. Unsteady simulations were performed using the NASA Common Research Model at a low-speed stall condition. The flowfield is characterized by large scale separation and a highly unsteady wake. An embedded high-resolution hexahedron block in the wake region of the grid has been used to improve resolution of the turbulence in unsteady Reynolds-Averaged Navier–Stokes and delayed detached eddy simulations. Model deformation occurring in the wind tunnel was incorporated into the computational setup. The computational fluid dynamics predictions match the force and wing pressure measurements from the wind tunnel very well. The mean and fluctuation data from the time-resolved particle image velocimetry measurements in the wake is matched by the delayed detached eddy simulations result, with the Reynolds-Averaged Navier–Stokes results showing stronger wake dissipation. Wake characteristics in delayed detached eddy simulations and Reynolds-Averaged Navier–Stokes have been compared using pressure spectra. The validity of the computational models has been evaluated as well.

Prediction and Measurement of the Common Research Model Wake at Stall Conditions

The development of the unsteady wake of transport aircraft under stall conditions and its impact on the empennage are challenging to predict, and the flow physics is far from being understood. In the European Strategic Wind Tunnels Improved Research Potential project, time-resolved particle image velocimetry measurements on the separated wake of the NASA Common Research Model were performed in the cryogenic European Transonic Windtunnel for flight Reynolds number conditions supplemented by aerodynamic measurements for a great variety of inflow conditions. In the time-resolved particle image velocimetry measurements, both low-speed stall (, ) and high-speed stall conditions (, ) were considered and the frequency resolution was as high as 1 kHz, enabling a sufficient characterization of the turbulent wake spectrum. For selected high- and low-speed stall conditions, hybrid Reynolds-averaged Navier–Stokes/large-eddy simulations were performed. In the present paper, the numerical results are discussed and compared to the experiments and first evaluations of the particle image velocimetry data are presented.

Reynolds Number Effect Investigation of Shock Wave on Supercritical Airfoil

The supercritical airfoil has been widely applied to large airplanes for sake of high aerodynamic efficiency. But at transonic speeds, the shock wave on upper surface of supercritical airfoil may induce boundary layer separation, which would change the aerodynamic characteristics. The shock characteristics such as location and intensity are sensitive to Reynolds number. In order to predict aerodynamic characteristics of supercritical airfoil exactly, the Reynolds number effects of shock wave must be investigated.The transonic flows over a typical supercritical airfoil CH were numerically simulated with two-dimensional Navier-Stokes equations, and the numerical method was validated with test results in ETW(European Transonic Windtunnel). The computation attack angles of CH airfoil varied from 0oto 8o, Mach numbers varied from 0.74 to 0.82 while Reynolds numbers varied from 3×106 to 50×106 per airfoil chord. It is obvious that shock location moves afterward and shock intensity strengthens as Reynolds number increasing. The similar curves of shock location and intensity is linear with logarithm of Reynolds number, so that the shock location and intensity at flight condition could be extrapolated from low Reynolds number.

Engineering Extrapolation to Flight Reynolds Number for Supercritical Airfoil Pressure Distribution Based on CFD Results

Pressure distribution of supercritical airfoil at flight Reynolds number could not be fully simulated except in cryogenic wind tunnel such as NTF (National Transonic Facility) and ETW (European Transonic Wind tunnel), which is costly and time resuming. This paper aimed to explore an engineering extrapolation to flight Reynolds number from low Reynolds number wind tunnel data for supercritical airfoil pressure distribution. However, the extrapolation method requiring plenty of data was investigated based on the CFD results for the reason of low cost and short period. Flows over a typical supercritical airfoil were numerically simulated by solving the two dimensional Navier-Stokes equations, with applications of ROE scheme spatial discretization and LU-SGS time march. Influence of computational grids convergence and turbulent models were investigated during the process of simulation. The supercritical airfoil pressure distribution were obtained with Reynolds numbers varied from 3.0×106 to 30×106 per airfoil chord, angles of attack from 0 degree to 6 degree and Mach numbers from 0.74 to 0.8. Simulated results indicated that weak shock existed on the upper surface of supercritical airfoil at cruise condition, that the shock location, shock strength and trailing edge pressure were dependent of Reynolds number, attack angles and Mach numbers. A similar parameter describing the Reynolds number effects factors was obtained by analyzing the relationship of shock wave location, shock front pressure and trailing edge pressure. Based on the similar parameter, airfoil pressure distribution at Reynolds number 30×106 was obtained by extrapolation. It was shown that extrapolated result compared well with simulated result at Reynolds number 30×106, implying that the engineering method was at least promising applying to the extrapolation of low Reynolds number wind tunnel data.

Engineering Extrapolation to Flight Reynolds Number for Supercritical Airfoil Pressure Distribution Based on CFD Results

Pressure distribution of supercritical airfoil at flight Reynolds number could not be fully simulated except in cryogenic wind tunnel such as NTF (National Transonic Facility) and ETW (European Transonic Wind tunnel), which is costly and time resuming. This paper aimed to explore an engineering extrapolation to flight Reynolds number from low Reynolds number wind tunnel data for supercritical airfoil pressure distribution. However, the extrapolation method requiring plenty of data was investigated based on the CFD results for the reason of low cost and short period. Flows over a typical supercritical airfoil were numerically simulated by solving the two dimensional Navier-Stokes equations, with applications of ROE scheme spatial discretization and LU-SGS time march. Influence of computational grids convergence and turbulent models were investigated during the process of simulation. The supercritical airfoil pressure distribution were obtained with Reynolds numbers varied from 3.0×106to 30×106per airfoil chord, angles of attack from 0 degree to 6 degree and Mach numbers from 0.74 to 0.8. Simulated results indicated that weak shock existed on the upper surface of supercritical airfoil at cruise condition, that the shock location, shock strength and trailing edge pressure were dependent of Reynolds number, attack angles and Mach numbers. A similar parameter describing the Reynolds number effects factors was obtained by analyzing the relationship of shock wave location, shock front pressure and trailing edge pressure. Based on the similar parameter, airfoil pressure distribution at Reynolds number 30×106was obtained by extrapolation. It was shown that extrapolated result compared well with simulated result at Reynolds number 30×106, implying that the engineering method was at least promising applying to the extrapolation of low Reynolds number wind tunnel data.

Simulation and model support correction for slotted wall transonic wind tunnels

Simulating the flow inside a wind tunnel can be useful for many purposes from planning experiments and design of a wind tunnel or components to CFD code validation or correction of wind tunnel measurements. This paper shows a CFD representation of a model plus support in a transonic wind tunnel flow and investigates its usefulness. Based on computational results, conclusions are drawn regarding support interference. The outline for a full support correction method is presented, including model support, mounting effects and buoyancy as well as other interferences due to non-uniform pressure distribution in the test section. The wind tunnel regarded is the European transonic windtunnel (ETW).

The use of computational fluid dynamics in TsAGI experimental works

The paper describes methods to calculate the flow around an aircraft model in a wind tunnel. It formulates special boundary conditions to achieve the necessary flow parameters in the control section, as well as to simulate some features of wind tunnel walls, for example, perforation. In order to accelerate the numerical solution convergence, a method for implicit smoothing is suggested allowing the calculation duration to be reduced several times. The cases of practical use of this methodology are given. It is shown that in the conditions of the TsAGI subsonic wind tunnel, it is possible to simulate the effect on the model from the elements of the structure that are missing in this tunnel, for example, the running track. A mathematical model of the European Transonic Windtunnel (ETW) with slotted walls is presented. It is shown that the flow in the reentry affects the main flow in the test section of this tunnel. Data on the effect of the model support in the TsAGI wind tunnel T-128 are given. The peniche height used in the half-model tests has been justified. The conclusion is made that the mathematical model of the wind tunnel is an obligatory part in experimental studies.

Reynolds Number Scaling Effects on the European High-Lift Configurations

Within the European High-Lift projects I and II, extensive experimental and numerical investigations on the aerodynamics of high-lift configurations have been carried out. The studies are conducted using the KH3Y windtunnel model DLR F11, which is representative of a wide-body commercial aircraft. A step-by-step complexity increase up to a complete high-lift configuration, including pylon, nacelle, and nacelle strake, is carried out to identify and separate the lift and drag behavior for the components of the high-lift system and their Reynolds number dependency. The wind-tunnel data have been gathered in the European Transonic Wind Tunnel in Cologne in three different test campaigns. To include the complete analysis of the stall behavior for the different configurations of a complex three-dimensional tapered high-lift wing with nacelles extends the scope of a single paper. Therefore, the focus of the present contribution is laid on the detailed analysis of Reynolds number scaling effects with respect to lift curves and drag polars. The combined variation of the dynamic pressure and the total temperature in the European Transonic Wind Tunnel allows covering a Reynolds number range from Re = 1,5 mil. up to deep cryogenic conditions with Re = 25 mil.. The final comparison of the four configurations reveals that the sequence of maximum lift values is maintained for the lowest and highest Reynolds number, although the increments between the configurations differ. The strongest increase in maximum lift is observed up to Re = 5 mil.. In the intermediate range, 5 mil. < Re < 15 mil, a highly nonlinear behavior of the maximum-lift values is observed, which requires further investigation.

Downward pointing winglet design and assessment within the M-DAW research project

AbstractThe design and performance analysis of a wing tip device proposed within the M-DAW project by ONERA is presented. A proto-design process is described and the device was thoroughly assessed (mainly with Reynolds-Averaged Navier-Stokes simulations). The process was further explained through wind-tunnel tests at both low speed and high speed in the pressurised and cryogenic European transonic wind tunnel in Cologne. The device is a downward pointing winglet designed for a retrofit scenario (the wing could be modified only within the 96% – 100% bounds of the span). It was designed to keep the wing root bending moment of the clean wing at cruise unchanged so that the aerodynamic gains are the net gains provided by the device that can be directly installed without structural modifications of the wing.

Advanced Design by Numerical Methods and Wind-Tunnel Verification Within European High-Lift Program

The design activity within the European 6th framework project EUROLIFT II is targeted towards an improvement of the take-off performance of a generic transport aircraft configuration by a redesign of the trailing edge flap. The involved partners applied different optimization strategies as well as different types of flow solvers in order to cover a wide range of possible approaches for aerodynamic design optimization. The optimization results obtained by the different partners have been cross-calculated in order to eliminate solver dependencies and to identify the best obtained design. The final selected design has been applied to the wind tunnel model and the test in the European Transonic Wind Tunnel (ETW) at high Reynolds number confirms the predicted improvements.

Dynamic qualification of the HIRENASD elastic wing model

Abstract The High Reynolds Number Aero-Structural Dynamics (HIRENASD) project is conducted by the Collaborative Research Center “Flow Modulation and Fluid-Structure Interaction at Airplane Wings (SFB 401)” at RWTH Aachen University. It is funded by the Deutsche Forschungsgemeinschaft (German Research Foundation, DFG) since 2004. Steady and unsteady state experiments with an elastic semispan wing model were performed under transonic and high Reynolds number conditions in the European Transonic Windtunnel (ETW) in Cologne, Germany. The main components of the complete windtunnel assembly were the wing model itself, a piezoelectric 6-components balance and a mechanical excitation mechanism able to excite the wing model at selected resonance frequencies. These components were designed, sized and assembled at the Department of Aerospace and Lightweight Structures at RWTH Aachen University. The dynamic qualification of the windtunnel assembly under wind-off condition was carried out in the department's test laboratory. Modal survey testing based on the hammer impact method was performed, as well as response decomposition analysis using enforced excitation by means of the mechanical excitation mechanism itself. Within the qualification, a test procedure enabling investigations of modal parameter changes under wind-off and wind-on conditions in a cryogenic windtunnel was also verified.

Investigation of a Half-Model High-Lift Configuration in a Wind Tunnel

An investigation is carried out for a half model high lift configuration inside the European Transonic Wind tunnel ETW. The influence of the wind tunnel walls and model installation is investigated and the numerical results are compared to measured data and free flight CFD results. The investigated model is a three element take-off configuration with full span slat and flap. A CFD solver for unstructured grids is used for the calculations. The computed results in the wind tunnel are in good agreement with uncorrected experimental data with maximum lift predicted at the same angle of attack. Corrected experimental and numerical tunnel data, however, deviate slightly from numerical results in free flight for which about 10% higher drag is predicted. In addition, the free flight maximum lift is predicted at a higher incidence. The lower drag in the in-tunnel results is due to a lower pressure at the leading edges of the slat and main wing close to the fuselage. This is a consequence of the current mounting of the wind tunnel model in the tunnel which causes a redistribution of the velocity field due to cross flow velocity components in the plane of symmetry of the half model. I. Introduction ALCULATING viscous fluid flows over high lift configurations is still a challenge in CFD. The difficulties in simulating these flows come from the complexity of both the geometry and flow physics. In particular, the multiple elements with small gaps give rise to multiple wakes, flow separation, laminar/turbulent transition, shock/boundary layer interaction etc., where many of these phenomena interact with each other. Since the fluid dynamics is dominated by viscous phenomena, only high-fidelity simulations based on the Navier-Stokes equations can provide the required accuracy to obtain realistic CFD solutions. The numerical simulation of the flow field around high lift configurations based on the Reynolds Averaged Navier-Stokes (RANS) equations has made significant progress during the last decade 1 . Until the beginning of the EUROLIFT project in early 2000, most of the European high lift activities had been devoted to two-dimensional computations 2 . The need for an extension to three dimensions as well as a state-of-the-art experimental database stimulated the launch of the EUROLIFT programme that was funded by the EC as part of the 5 th European framework program. A close coupling and harmonization between experimental and numerical activities was attempted in the project. CFD was brought into a more daily use in EUROLIFT and introduced in three dimensions. Mainly through hybrid Navier-Stokes technology, it has become possible to compute viscous flows about take-off and landing configurations within a reasonable time frame and with sufficient reliability. The work carried out in the EUROLIFT project has resulted in several publications, an overview is given in references 3-6 . EUROLIFT II 7 is a European High Lift Programme in the 6 th European framework following the work initialized in EUROLIFT. The numerical and experimental investigations in EUROLIFT left many important questions unanswered being pursued in EUROLIFT II. The investigation described here is concerned with the installation effects inside the cryogenic wind tunnel ETW (European Transonic Wind tunnel) in which experiments at high Reynolds numbers are conducted. In particular, the influence of the wind tunnel walls and model installation is investigated by conducting CFD calculations inside the wind tunnel in comparisons with measurements and with computational results from free flight calculations. Numerous CFD calculations in free flight have been carried out with comparisons to experimental results from ETW in which reasonable agreement is reached in lift between numerical results and experiments. The maximum lift, however, is often predicted too late at a higher angle and, in particular, the drag is almost always over-predicted. This investigation has been conducted to establish if this

Problems in practical application of high resolution numerical schemes to internal and external tasks of aerodynamics

AbstractThis report presents numerical method for solution of aerodynamic tasks on the basis of Reynolds Averaged Navier‐Stokes Equations. The method is applied to simulation of flow around aircraft (including engine‐airframe interference) and to description of flowfield in the wind tunnel (WT) with aircraft model inside. To satisfy growing practical requirements both to accuracy of numerical results and to calculation time, zonal approach is proposed: flow regions with small scales of physical phenomena may be described using implicit scheme, and the rest of the flowfield – using explicit scheme. An example of modeling flow in the test section of European Transonic Windtunnel is shown. Unsteady phenomena around WT slotted walls were obtained. Calculations of noise suppressing nozzles with different geometry are considered. Numerical simulation has allowed to improve the nozzle geometry and to provide essential noise suppression. All presented examples confirm the need in further development of effective and accurate algorithms for practical tasks solution. (© 2008 WILEY‐VCH Verlag GmbH &amp; Co. KGaA, Weinheim)

A numeric far field model for support interference studies in a slotted wall wind tunnel (European Transonic Windtunnel)

This article presents a simplified computational fluid dynamics (CFD) model of the European Transonic Windtunnel test section that represents the model support in a far field and is specifically designed for the investigation of model support effects in the transonic regime. Two parametric studies were carried out in order to determine geometric parameters of the adapted support geometry and in order to enable the simulation of actual wind tunnel experiments, by correlating the wind tunnel reference Mach number with that at the far field in the CFD computation and the effective Mach number at the model location. A comparison was done using measured calibration data from the wind tunnel.

B305 大型低温風洞で用いる感温塗料の開発

A new Temperature-Sensitive Paint (TSP) technique for boundary-layer transition detection in a production-type large cryogenic wind tunnel is present. The formulation of Ruthenium complex TSP system has been optimized in terms of luminescence intensity and robustness. The optimum dye-binder-solvent combination has been determined through systematic sample tests. A new binder has been introduced and the resulting coating was found free from cracking at cryogenic temperatures. This is contrary to the silicone-based cryogenic TSPs that are subject to micro cracks at reduced temperatures. The new TSP can meet the root-mean-square roughness requirement less than 0.15 ?m. We applied this TSP in the JAXA 0.1-m Transonic Cryogenic Wind Tunnel, and then European Transonic Windtunnel (ETW). A variety of transition patterns have been successfully visualized.

ETW: simulations of true flight behaviour

The capabilities of the European Transonic Windtunnel ETW to simulate flight conditions by testing at cryogenic temperatures are presented. The achieved flow quality and tunnel control form the basis for an excellent data repeatability and a high confidence level in the deliverables. Unique instrumentation has been developed to provide additional information about the test object and the surrounding flow field during tunnel entries. Typical samples of Reynolds number effects are shown. Productivity aspects are considered on the basis of recently performed test scenarios.