European Transonic Windtunnel Research Articles

Aeroacoustic testing on a full aircraft model at high Reynolds numbers in the European Transonic Windtunnel

This paper presents an end-to-end approach for the assessment of pressurized and cryogenic wind tunnel measurements of an EMBRAER scaled full model close to real-world Reynolds numbers. The choice of microphones, measurement parameters, the design of the array, and the selection of flow parameters are discussed. Different wind tunnel conditions are proposed which allow separating the influence of the Reynolds number from the Mach number, as well as the influence of slotted and closed test sections. The paper provides three-dimensional beamforming results with CLEAN-SC deconvolution, the selection of regions of interest, and the corresponding source spectra. The results suggest that slotted test sections have little influence on the beamforming results compared to closed test sections and that the Reynolds number has a profound, non-linear impact on the aeroacoustic emission that lessens with increasing Reynolds number. Further, sources show a non-linear Mach number dependency at constant Reynolds number but are self-similar in the observed Mach number range. The findings suggest that it is possible to study real-world phenomena on small-scale full models at real-world Reynolds numbers, which enable further investigations in the future such as the directivity of sources.

Mach and Reynolds number effects on transonic buffet on the XRF-1 transport aircraft wing at flight Reynolds number

This work provides an overview of aerodynamic data acquired in the European Transonic Windtunnel using an XRF-1 transport aircraft configuration both at cruise conditions and at the edges of the flight envelope. The goals and design of the wind tunnel test were described, highlighting the use of the cryogenic wind tunnel’s capability to isolate the effects of M_{infty }, {text{Re}}_{infty } and the dynamic pressure q/E. The resulting dataset includes an aerodynamic baseline characterization of the full span model with vertical and horizontal tailplanes and without engine nacelles. The effects of different inflow conditions were studied using data from continuous polars, evaluating the changes in aeroelastic deformation which are proportional to q/E and the influence of M_{infty } and {text{Re}}_{infty } on the shock position. Off-design data was analyzed at the lowest and highest measured Mach numbers of 0.84 and 0.90, respectively. Wing lower surface flow and underside shock motion was analyzed at negative angles of attack using {{c}_{text{p}}} distribution and unsteady pressure transducer fluctuation data, identifying significant upstream displacement of the shock close to the leading edge. Wing upper-side flow and the shock motion near buffet onset and beyond was analyzed using unsteady pressure data from point transducers and unsteady pressure-sensitive paint (PSP) measurements. Buffet occurs at lower angles of attack at high Mach number, and without clearly defined lift break. Spectral contents at the acquired data points in the buffet range suggest broadband fluctuations at Strouhal numbers between 0.2 and 0.6, which is consistent with recent literature. The spanwise shock propagation velocities were determined independently via analysis of unsteady PSP and pressure transducers to be in the range between u_{text{s}} / u_{infty } = 0.24 and 0.32, which is similarly in line with published datasets using other swept wing aircraft models.

Prediction of wing buffet pressure loads using a convolutional and recurrent neural network framework

AbstractIn the present study, a hybrid deep learning reduced-order model (ROM) is applied for the prediction of wing buffet pressure distributions on a civil aircraft configuration. The hybrid model is compound of a convolutional variational neural network autoencoder (CNN-VAR-AE) and a long short-term memory (LSTM) neural network. The CNN-VAR-AE is used for the reduction of the high-dimensional flow field data, whereas the LSTM is applied to predict the temporal evolution of the pressure distributions. For training the neural network, experimental buffet data obtained by unsteady pressure sensitive paint measurement (iPSP), is applied. As a test case, the Airbus XRF-1 configuration is selected, considering two different experimental setups. The first setup is defined by a wind tunnel model with a clean wing, whereas the second setup includes an ultra high bypass ratio engine nacelle on each wing. Both configurations have been tested in the European Transonic Windtunnel, considering several transonic buffet conditions. Finalizing the training of the hybrid neural networks, the trained models are applied for the prediction of buffet flow conditions which are not included in the training data set. A comparison of the experimental results and the pressure distributions predicted by the hybrid ROMs indicate a precise prediction performance. Considering both aircraft configurations, the main buffet flow features are captured by the hybrid ROMs.

Time-resolved pressure-sensitive paint measurements for cryogenic wind tunnel tests

A first time-resolved pressure-sensitive paint (PSP) test campaign at the European Transonic Wind Tunnel (ETW) was conducted within the research initiative “Unsteady flow and interaction phenomena at High Speed Stall conditions”. One of the objectives of this wind tunnel campaign was to resolve time-series of surface pressure distributions caused by complex 3-D buffet phenomena on a full-span airplane model XRF-1 transport aircraft configuration. At higher angle-of-attack and high Mach number, pressure fluctuations with a frequency of several hundred Hertz are expected to occur on the main wing and the horizontal tail plane (HTP) caused by the buffet effect. To capture the expected buffet phenomena by PSP, the German Aerospace Center developed a time-resolved PSP measurement and data acquisition system as well as a post-processing method for measurements in ETW. The measurements were conducted on the main wing and HTP simultaneously, with a camera frame rate of up to 2 kHz. The transonic buffet phenomena were observed at the flight relevant Reynolds number Re = 12.9 × 106 and Re = 25.0 × 106. The time-varying surface pressure distribution on the model was successfully captured by PSP. The time-series and spectra of both PSP and pressure transducer data match very well.

Research Initiative for Numerical and Experimental Studies on High-Speed Stall of Civil Aircraft

The aerodynamics of transport aircraft at the borders of the flight envelope is characterized by complex interactions and nonlinear, unsteady flow phenomena. The underlying physical mechanisms are not fully understood, and a prediction of aerodynamic properties, transient loads, and aeroelastic behavior is a major challenge that requires the use of sophisticated numerical models and dedicated experiments. A new research initiative investigates the three-dimensional buffet mechanisms at high-speed stall, the interaction of the wing wake with the empennage, and the influence of ultra-high-bypass-ratio-type engines on these phenomena. It comprises unsteady pressure-sensitive paint surface pressure and time-resolved particle image velocimetry flowfield measurements carried out in the cryogenic European Transonic Windtunnel. The experiments are conducted at different Mach numbers and cover a broad Reynolds number regime up to using the generic Airbus XRF1 configuration as a reference. This paper explains the rationale of the research initiative, introduces the planned test entries and test conditions, and presents selected results of the first measurement campaign and associated numerical analyses by means of the computational fluid dynamics code TAU, including usage of locally scale resolving simulations. An analysis of numerical and experimental results is presented for a high incidence transonic condition that serves to discuss the topology of wing root flow separation, the evolution of the associated wake, and its interaction with the empennage.

Prediction of transonic wing buffet pressure based on deep learning

In the present study, a deep learning approach based on a long short-term memory (LSTM) neural network is applied for the prediction of transonic wing buffet pressure. In particular, fluctuations in surface pressure over a certain time period as measured by a piezoresistive pressure sensor, are considered. As a test case, the generic XRF-1 aircraft configuration developed by Airbus is used. The XRF-1 configuration has been investigated at different transonic buffet conditions in the European Transonic Wind tunnel (ETW). During the ETW test campaign, sensor data has been obtained at different local span—and chordwise positions on the lower and upper surface of the wing and the horizontal tail plane. For the training of the neural network, a buffet flow condition with a fixed angle of attack alpha and a fixed sensor position on the upper wing surface is considered. Subsequent, the trained network is applied towards different angles of attack and sensor positions considering the flow condition applied for training the network. As a final step, the trained LSTM neural network is used for the prediction of pressure data at a flow condition different from the flow condition considered for training. By comparing the results of the wind tunnel experiment with the results obtained by the neural network, a good agreement is indicated.

Investigation of Reynolds Number Effects on Aerodynamic Characteristics of a Transport Aircraft

The scale difference between the real flight vehicle and the experimental model results in the Reynolds number effect, which makes it unreliable to predict the aerodynamic characteristics of flight vehicles by wind tunnel testing. To understand the mechanism of Reynolds number effects on the aerodynamic characteristics of the supercritical wing that is commonly used in transport aircraft in more detail, surface pressure wind tunnel tests of a transport aircraft reference model with a wing-body configuration were conducted in the European Transonic Windtunnel (ETW) at different Reynolds numbers. There are 495 pressure taps in total equipped on the surface of the test model with the Mach numbers ranging from 0.6 to 0.86 and Reynolds number varying from 3.3 × 106 to 35 × 106. In addition, an in-house developed CFD tool that has been validated by extensive experimental data was used to correct the wing deformation effect of the test model and achieve detailed flow structures. The results show that the Reynolds number has a significant impact on the boundary layer displacement thickness, surface pressure distribution, shock wave position, and overall aerodynamic force coefficients of the transport aircraft in the presence of shock wave and the induced boundary layer separation. The wind tunnel data combined with flow fields achieved from CFD show that the essence of the Reynolds number effect on the aerodynamic characteristics of transport aircraft is the difference of boundary layer development, shock wave/boundary layer interaction, and induced flow separation at different Reynolds numbers.

A CSM-CFD methodology applied to the design of a cryogenic WT model

The Aero-structural coupled analysis presented in this work represents an important part of the PRODIGE project, whose main purpose is the prediction of aerodynamics and hinge moment loads of the aileron of a scale model of business jet at transonic Mach numbers and flight Reynolds number. More in details, this methodology supports the design activities of the cryogenic wind-tunnel (WT) model to be tested in the European Transonic Wind tunnel (ETW). The need to implement such a kind of analysis comes from the requirement to take into account the “pure aeroelastic effect” during experimental tests and need to understand the effect of the WT model flexibility on the control surface load prediction, prior to and in parallel with the balance design. CFD evaluations are performed on the whole model, while a static Fluid-Structure-Interface (FSI) procedure is set up and, only, applied to the wing and aileron of the model. The FSI cycle is made of two blocks: a CFD evaluation part and a CSM evaluation part. Both modules are coupled within an iterative loop, where information are exchanged between them. Two angles of incidence are taken into account, i.e. the lowest value (α=-3°) and the highest one (α=6°). Results will be compared in terms of global aerodynamic forces and pressure coefficient distributions. Moreover the new deformed shapes derived from the application of both approaches will be compared with the original one.

Reynolds number effects of supercritical airfoil with mini-TED in transonic flow

In order to accurately investigate the Reynolds number influence of supercritical airfoil with miniature trailing-edge devices (mini-TED) under transonic conditions. The supercritical airfoil pressure measurement test of high Reynolds number was conducted in the European transonic wind tunnel. The reliability of the high Reynolds number numerical calculation method was validated by the wind tunnel test results. The calculation results are in good agreement with the wind tunnel test results, indicating that the high Reynolds number calculation accuracy could meets the investigation requirement. The transonic flow of supercritical airfoil with mini-TED in the range of Reynolds number 3×106∼50×106 and Mach number 0.76 was simulated using the validated calculation method. Due to the existence of mini-TED, the Reynolds number influence self-adjustment range of the supercritical airfoil disappeared, while the influence magnitude and sensitivity reduced. At the same time, the supercritical airfoil’s linear influence of Reynolds number effects has been changed, and the Reynolds number prediction method based on pressure distribution is no longer applicable. After considering the influence of Reynolds number, the detailed design of supercritical wing with mini-TED could be conducted to improve its reliability and economy in the application of advanced aircraft.

Evaluation of Cryogenic Force Balance Calibration Methodologies Relative to Wind Tunnel Results

Independent tests of the NASA Common Research Model at NASA’s National Transonic Facility and the European Transonic Windtunnel reveal differences at low operating temperatures and high Reynolds numbers that warrant further investigation. Since each facility used their own wind tunnel balance for their tunnel entry, one suggestion for the differences was the temperature compensation methodology developed and applied for each balance. This hypothesis is explored through simulation and experiment. Independent calibrations of NASA’s NTF-118A balance at NASA Langley Research Center and European Transonic Windtunnel reveal differences in the thermal compensation of the normal force and pitching moment primary sensitivities with temperature, while the axial force primary sensitivities are in good agreement. The application of the balance calibrations performed at NASA and European Transonic Windtunnel to the prior wind tunnel data suggests that the thermal compensation differences are an order of magnitude lower than the differences observed between the wind tunnel aerodynamic coefficients. Thus, the balance temperature compensation methodologies used by NASA and European Transonic Windtunnel are not a major contributor to the wind tunnel differences.

Wind Tunnel Support System Influence on NASA Common Research Model at Low-Speed Conditions

Computational studies were conducted in order to gain an understanding and to quantify the impact of the model support system and aeroelastic wing deformation on the Common Research Model at subsonic inflow conditions at . Comparisons between results obtained in the European Transonic Windtunnel and from simulations using TAU reveal the extent of the modeling error introduced by neglecting the support structure. RANS simulations with and without the support system show that it has a near-field effect on the model’s rear fuselage, significantly altering the pressure field and thereby impacting the force coefficients, most significantly drag and pitching moment. Aeroelastic wing deformation was also included in the investigation, showing non-negligible impact on the forces. The inclusion of the support system and wing deformation, as well as the use of advanced turbulence models, was shown to promise considerable error reduction potential in comparative studies between wind tunnel and simulation data.

ESWIRP: European Strategic Wind tunnels Improved Research Potential program overview

“European Strategic Wind tunnel Improved Research Potential” ESWIRP was a project in the EU 7th Framework Program (FP7—Grant agreement no: FP7-227816), which was aiming at improving the performance capabilities of three strategic wind tunnels in Europe, by strengthening the cooperation between these wind tunnels in a new consortium. The research consortium members are Office National d’Etudes et de Recherches Aérospatiales (operating the S1MA as its largest sonic wind tunnel), German–Dutch wind tunnels [operating the large low-speed facility (LLF) as its largest low-speed wind tunnel], and European transonic wind tunnel (ETW) (operating its cryogenic wind tunnel). Together, these wind tunnels cover a wide range of experimental capabilities of relevance to civil aviation and aeronautical research in general. The project started in October 2009 for a period of 5 years. The European financial contribution was €7.2 million. The project consist of two major parts: (1) improvements to the testing infrastructure; and (2) the provision of wind tunnel access to research groups which do not usually have the means to access such large-scale test facilities. These topics also involved public dissemination and information activities. Although the tunnels covered in this project are of a complementary nature, the infrastructure activities were joined together, by a common representation of, and approach to, the tunnel performance characteristics. To this end, a generic model of a virtual wind tunnel was developed, enabling operators to assess the effect of the control parameters upon the testing conditions. The final aim of all participants was to provide the user community with an improved set of capabilities to test their innovative ideas. To provide better access to these three major wind tunnels, mainly research groups from European universities were contacted. The approach taken has included maximum transparency of the process and support of the researchers by the organizations responsible for the tunnels. In addition, when possible, we encouraged research groups to work together, to obtain the full benefit of economies of scale in research projects. ESWIRP responded to the targeted approach of the Integrating Activities of the FP7 Capacities Work Program: Networking activities, essentially focused around four topics: (1) organization of information campaigns, lectures and workshops to disseminate knowledge between the partners and future users. (2) Opening of a website for the consultation of wind tunnel standards. (3) Exchange of personnel between the Consortium partners, to foster the spread of good practices and the exchange of technical know-how. (4) Joint development of a reference wind tunnel parameter database. Trans-national access and/or testing services: after the call for proposals by the facility providers, groups of researchers had the opportunity to benefit from free wind tunnel services, including technical assistance to support the corresponding scientific research team(s). Joint research activities, innovative modeling of wind tunnels has helped designers to make better decisions, before the implementation of any novel hardware. Based on a generic finite volume type of modeling approach, the operational behaviour and the time-dependent flow quantities around the wind tunnel circuit have been simulated. Such model can be used, for example, to conceive and tune particular sub-system control laws, or explore the implications of hardware changes. Based on a common approach, dedicated models were developed for S1MA, LLF and ETW. The infrastructure improvements targeted the capability to obtain unsteady test data (with high accuracy) in the ETW, to improve the capability to simulate aircraft behaviour in ground effect in the LLF, and to establish a reliable closed-loop control of test Mach number in the S1MA wind tunnel. European Strategic Wind tunnels Improved Research Potential Program is European support for strategic wind tunnels, key research infrastructures in the development process of current and future aircraft. This is the first time that European authorities have given such support.

Using wing model deformation for improvement of CFD results of ESWIRP project

The NASA Common Research Model has been used for evaluation of the effect of the deformation of the model on aerodynamic forces and local flowfield. Large differences were observed during comparison of the results of the rigid model and experimental data obtained during ESWIRP project at European Transonic Windtunnel. On-line mesh deformation using 20 modes in total was used during coupled CFD simulations to improve the accuracy of the numerical results. The results from these simulations showed that adding the wing deformation into the computational fluid dynamics simulation improved the accuracy of obtained aerodynamic coefficients and also the local flowfield. The results from the coupled CFD simulation with on-line mesh deformation of the semispan model are the focus of this paper.

Investigation of viscosity influence on transonic flutter

Abstract The paper is devoted to computational study of transonic flutter when the viscosity can influence significantly on dynamic aeroelasticity characteristics of aircraft. The work has been done in the direction of development of methodology and software, which are used in ARGON (TsAGI) system for multidisciplinary analysis and optimization in airplane design. Considerable attention is paid to the validation of the proposed software. Experimental results of the NASA Common Research Model in the European Transonic Wind Tunnel (ETW) are used for comparison with computations. The results of flutter analysis are presented for the passenger middle range airplane with the high aspect ratio wing and the engine under the wing. Comparisons of aeroelasticity characteristics in transonic flow are carried out for cases of a set of Mach and Reynolds numbers. The computational results presented in the paper show the essential influence of transonic features on flutter characteristics.

Wind-Tunnel and CFD Investigations Focused on Transition and Performance Predictions of Laminar Wings

This paper is a continuation of the paper recently published by Hue et al. (“Experimental and Numerical Methods for Transition and Drag Predictions of Laminar Airfoils,” AIAA Journal, Vol. 53, No. 9, Sept. 2015, pp. 2694–2712) that focused on transition and drag predictions of laminar airfoils. The extension of such studies to three-dimensional configurations representative of modern civil aircraft is a further step toward the implementation of natural laminar-flow technologies. The present work, therefore, focuses on validating the laminar design of a low-swept wing for business jet applications. In 2015, an experimental campaign was carried out in the European Transonic Windtunnel, and included laminar/turbulent transition measurements with temperature-sensitive paint at Mach and Reynolds numbers typical of cruise flight conditions. Subsequently, fluid dynamics computations were performed on this aircraft geometry either with a Reynolds-averaged Navier–Stokes solver using both Tollmien–Schlichting and crossflow transition criteria, or with a boundary-layer code using either database methods for transition location or exact stability analyses. In this paper, experimental and numerical transition predictions are compared for three representative cases corresponding to different angles of attack. The agreement that is achieved is satisfactory, and extended regions of laminar flow are observed on the wing at cruise lift levels. In these conditions, the drag reduction can account for 10–15% of the aircraft drag.

High-speed PIV applied to wake of NASA CRM model at high Re-number sub- and transonic stall conditions

Within the framework of the EU project ESWIRP, the particle image velocimetry (PIV) using high-speed camera and laser has been used to measure the turbulent flow in the wake of a stalled aircraft wing. The measurements took place on the common research model provided by NASA in the pressurized cryogenic European Transonic Wind tunnel. A specific cryoPIV system has been used and adapted for using high-speed PIV components under the cryogenic conditions of the wind tunnel facility. First results are presented comprising transonic and subsonic stall conditions at realistic flight Reynolds numbers of 11.6 and 30 million, respectively.

Development of Deformed Computer-Aided Design Geometries for the Sixth Drag Prediction Workshop

A method for modifying the shape of digital geometry representations used in computer-aided design applications according to experimental pointwise deflection data is described. In the pilot study presented here, the method is applied to the computer-aided design geometry of an aircraft wing that is deformed using measured deflections obtained from the associated wing model during a wind-tunnel test. The method provides various advantages in computational fluid dynamics and other applications. Numerical grids generated on the deformed geometry yield an improved correlation of computational fluid dynamics results to wind-tunnel data without the need for aeroelastic computations, which require coupling to a structural model. Additionally, the method is beneficial in aerodynamic optimization, where the optimized shape is usually given in the form of a computational fluid dynamics surface mesh or nodal deflections that, for further processing or manufacturing purposes, need to be transferred into a computer-aided design geometry description. The paper presents an application to NASA’s Common Research Model transport aircraft configuration, where deformation measurements from a test campaign in the European Transonic Wind Tunnel are processed to provide deformed computer-aided design geometries for the Sixth AIAA Computational Fluid Dynamics Drag Prediction Workshop.

Comparison of the NASA Common Research Model European Transonic Wind Tunnel test data to NASA National Transonic Facility test data

Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility and the European Transonic Wind Tunnel. Data have been obtained at chord Reynolds numbers of 5, 19.8 and 30 million for the wing/body/tail = 0° incidence configuration in the National Transonic Facility and in the European Transonic Wind Tunnel. Force and moment, surface pressure, wing bending and twist, and surface flow visualization data were obtained in both facilities but only the force and moment, and surface pressure data are presented herein.

Experimental Investigations on Common Research Model at ONERA-S1MA–Drag Prediction Workshop Numerical Results

This paper aims to present some of the experimental and numerical results obtained with the NASA–Boeing Common Research Model at ONERA. The wind tunnel model used in the present study is the ONERA Large Reference Model (1/16.835), which has the same geometry as the Common Resarch Model considered in the latest AIAA Drag Prediction Workshops. Experimental data have been collected from the ONERA-S1MA wind tunnel at Mach numbers between 0.30 and 0.95 and a mean aerodynamic chord Reynolds number of 5 million for four different configurations: wing/body only and wing/body with horizontal or vertical tails or both. Force and moment and pressure and surface flow measurements have been performed. Concerning the numerical study, all the Reynolds-averaged Navier–Stokes computations have been completed with the structured solver elsA, and the Spalart–Allmaras and with shear stress transport turbulence models have been used in addition to the quadratic constitutive relation. In this paper, configuration effects (increments due to horizontal and/or vertical tails) are assessed both numerically and experimentally for several Mach numbers and angles of attack. The delicate issue of flow separation at the wing/body junction is also addressed with the support of oil flow visualizations. elsA and S1MA skin pressure distributions are presented; the agreement is satisfactory except for some outboard wing sections at high lift levels. Finally, comparisons of drag and moment values including computational fluid dynamics and test data from different wind tunnels are proposed (S1MA, NASA Ames Research Center and National Transonic Facility, and European Transonic Wind Tunnel).

Summary of Case 5 from Sixth Drag Prediction Workshop: Coupled Aerostructural Simulation

A summary of participants’ results for the coupled aerostructural simulation test case from the Sixth AIAA Computational Fluid Dynamics Drag Prediction Workshop held in Washington, DC, 16–17 June 2016, is presented. The test case focuses on a static aeroelastic simulation of NASA’s Common Research Model civil transport aircraft in the wing/body configuration (i.e., without engines, pylons, and empennage). Flow conditions include the design point lift coefficient at cruise flight. Aeroelastic simulations are performed on the “medium” baseline grid, coupled to a computational structural analysis, using either an existing finite element model or modal data. Experimental validation data for this test case are available from a wind-tunnel test campaign performed at the European Transonic Wind Tunnel in Cologne, Germany.