Ralf Rudnik

EUROLIFT Test Case Description for the 2nd High Lift Prediction Workshop

The paper describes the experimental evidence for the DLR-F11 high lift configuration to be used within the context of the 2 nd phase of the AIAA High Lift Prediction Workshop. The model geometry is representative for a wide-body commercial aircraft. For the present purpose a wing/body combination is considered with a continuous slat and flap system in landing setting. Slat and flap are intersecting with the fuselage in order to suppress side edge interference effects and their aerodynamic impact on maximum lift. A CAD model in various degrees of detail has been refurbished, serving as the common geometrical basis for the scheduled CFD investigations. Experimental data of the European project EUROLIFT for low and high Reynolds number conditions have been made available, making use of the same wind tunnel model. The data for low Reynolds numbers have been gathered in the Low Speed Wind Tunnel of Airbus in Bremen, B-LSWT, Germany, while the high Reynoldsnumber data have been measured in the European Transonic Windtunnel, ETW, under cryogenic conditions. The Reynolds numbers between both datasets differ by an order of magnitude. In addition to force and moment data, which are available from both wind tunnel tests, a comprehensive validation database is available of the tests in the B-LSWT. The experimental data comprise oil flow pictures, transition information by hotfilms and infrared thermography, as well as PIV velocity data in various locations of the F11 configuration for a sample of angles of attack up to and beyond maximum lift. The main features of the experimental evidence are analyzed, comparing pressures and forces for low and high Reynolds number conditions. Examples of the oil flow pictures, transition information, and off-body velocity data are presented and briefly discussed.

Three-dimensional Navier-Stokes simulations for transport aircraft high-lift configurations

Computations of lift and drag polars for a transport aircraft wing/fuselage high-lift cone guration using the MEGAFLOWcodesystemarecarriedoutandcomparedtowind-tunnelexperiments.Themainemphasisislaidon acomparisonoftheblock-structuredandtheunstructuredcodemodulesforsuchtypeofapplication.FortheblockstructuredFLOWercodeincombinationwith a k‐! turbulencemodel,thenumericalresultsarein good agreement with the available experimental data in the linear CL range. Beyond 15-deg incidence, a strong separation near the e ap cut-out is simulated, leading to an underprediction of total lift near CL; max compared to the experimental data. In contrast to this, the results of the unstructured TAU code utilizing the Spalart ‐Allmaras turbulence model are characterized by a nearly constant lift overestimation up to maximum lift without the aforementioned separation tendency at moderate incidences. The lift overprediction in the unstructured results is attributed to the main wing and the slat upperside suction peaks, which are higher resolved by the unstructured grid. Neither code reproduces the lift breakdown beyond CL; max according to the experiments. The use of preconditioning in conjunction with theFLOWercodeshowsonly minorimprovement of theaccuracy,but considerabledeterioration of the convergence properties, requiring improvements for routine use. Further studies will focus on the ine uence of geometry simplie cations at the wing root in the theoretical models and its impact on the experimental evidence.

DLR Contribution to the First High Lift Prediction Workshop

The paper describes numerical studies by DLR devoted to the computation of the DLR F-11 high lift configuration as part of the 2nd phase of the AIAA High Lift Prediction Workshop. The model geometry is representative for a wide-body commercial aircraft with a classical three element high lift system at the wing leading and trailing edge in a landing setting. Being a follow-up of the NASA Trapezoidal wing of the 1st phase of the workshop, the F-11 wing/body wind tunnel model features only a modest overall complexity increase in the basic set-up. Yet, the most complex geometry representation includes model details, such as slat tracks, flap track fairings, and pressure tube bundles. According to the available experimental evidence, low and high Reynolds numbers have been investigated. In addition to force, moment, and pressure distributions, surface streamlines have been analyzed. The studies have been carried out using the DLR TAU code in conjunction with the Spalart-Allmaras turbulence model in the original formulation. DLR contributed to the grid generation activities by providing a set of hexahedral-based hybrid unstructured grids for all complexity stages and both Reynolds-numbers. For the most simplified configuration, a family of three grids has been generated to study grid resolution aspects. In general, a good agreement to the experimental data is obtained. It turns out that for the present wind tunnel configuration, the inclusion geometry details of the wind tunnel model, such as slat tracks and even pressure tube bundles, is essential for the agreement to the experimental evidence and the prediction of maximum lift.

Experimental Investigation of the DLR-F6 Transport Configuration in the National Transonic Facility

An experimental aerodynamic investigation of the DLR (German Aerospace Center) F6 generic transport configuration has been conducted in the NASA NTF (National Transonic Facility) for CFD validation within the framework of the AIAA Drag Prediction Workshop. Force and moment, surface pressure, model deformation, and surface flow visualization data have been obtained at Reynolds numbers of both 3 million and 5 million. Flow-through nacelles and a side-of-body fairing were also investigated on this wing-body configuration. Reynolds number effects on trailing edge separation have been assessed, and the effectiveness of the side-of-body fairing in eliminating a known region of separated flow has been determined. Data obtained at a Reynolds number of 3 million are presented together for comparison with data from a previous wind tunnel investigation in the ONERA S2MA facility. New surface flow visualization capabilities have also been successfully explored and demonstrated in the NTF for the high pressure and moderately low temperature conditions required in this investigation. Images detailing wing surface flow characteristics are presented.

Numerical Simulation of Engine Airframe Integration for High-Bypass Engines

Abstract Due to a trend towards very high-bypass ratio engines and a corresponding close coupling of engine and airframe, the minimization of adverse interference effects is an important aspect in aircraft design. Investigations of engine/airframe integration have been carried out within a long-term collaborative European research initiative, starting in 1990 with the programs DUPRIN I, DUPRIN II to the current ENIFAIR and AIRDATA projects. Based on some selected results the contribution highlights major outcomes of the numerical activities accompanying the experimental studies in the aforementioned programs. After a brief introduction to the basic aerodynamic phenomena of engine/airframe interference and the numerical methods in use, the capabilities of the theoretical approach are demonstrated for three aspects: The influence of increasing engine size on the aerodynamic interference is outlined by simulating turbine powered engine simulators (TPS) of different bypass ratio on the ALVAST narrow body wing/fuselage model. Second, the influence of position variations is demonstrated for different engine concepts, representing the major design parameter for influencing engine/airframe interference. Finally, the jet influence is stressed by comparing numerical results for different thrust conditions. The investigations show, that the lift loss, caused by the mounting of engines, is proportional to the engine size. An upstream movement of the engine position alleviates the lift loss, whereas a vertical movement does not have a significant influence. Especially for the VHBR and UHBR concepts the incorporation of the engine jet is essential to assess the aerodynamic interference. In general validated numerical methods are capable to simulate the dominant features of engine/airframe integration.

Euler computation of the nearfield wake vortex of an aircraft in take-off configuration

Abstract Numerical simulations based on the three-dimensional Euler equations are used to investigate the predictive capability of an Euler code for calculations of the nearfield wake of a narrow-body airliner wind tunnel model in take-off configuration up to a half span behind the wing tip trailing edge. Simulation results on both structured and unstructured grids are presented. The results on the block-structured grid were obtained within the scope of the EU-project EUROWAKE. The simulation quality of the vortex formation and spatial development is analysed by comparison to wind tunnel measurements of the spanwise lift distribution available from the EU-project DUPRIN II and to experimental PIV data available from the EU-project EUROWAKE.

Numerical Investigation of Engine Effects on a Transport Aircraft with Circulation Control

The scope of this paper is to illustrate the installation effects of a turboprop engine on a high-lift configuration of a short takeoff and landing aircraft with circulation control. In addition to the influence on the wing performance, the impact on the longitudinal static stability of the aircraft is also investigated. Furthermore, critical failure cases, namely one engine inoperative, as well as an asymmetric circulation control failure, are assessed. Therefore, steady computational-fluid-dynamics calculations based on the Reynolds-averaged Navier–Stokes equations were performed. The propeller is modeled with an actuator-disk approach. The results show strong potential of increasing lift by synergy effects between circulation control and propeller slipstream. However, the longitudinal stability and controllability are adversely affected. Regarding the case of one engine inoperative, the resulting yawing moments are twice as high as the actual yawing moments from the asymmetric thrust and are therefore ...

Experimental Investigation of the Wing-Body Juncture Flow on the DLR-F6 Configuration in the ONERA S2MA Facility

An experimental aerodynamic investigation of the DLR F6 generic transport aircraft configuration has been conducted in the S2MA wind tunnel facility of ONERA in Modane, France, in a collaboration between DLR and ONERA in November 2008. The test with the DLR-F6 configuration is a follow-on study of tests in this tunnel in the 1990s, as well as of recent tests with the same configuration in NASA NTF facility in 2007, which has been carried out by NASA and DLR as a contribution to the AIAA Drag Prediction Workshop activities. While the test campaign in the NASA NTF covered the wing/body, the wing/body/fairing, and the wing/body/nacelle/pylon configuration for Reynolds-numbers of Re=3 x 10 6 and 5 x 10 6 , the present experimental investigation is limited to the wing/body configuration with and without a side-of-body fairing for a Reynolds-number of Re=3 x 10 6 . The main purpose of the test is to address specific issues and findings from the previous NASA NTF test, as well as to provide more detailed and additional flow field information of the wing/body juncture flow for these two configurations with different measurement techniques. The contribution is confined to the analysis of results of the ONERA S2MA wind tunnel. The measurements of force and moment data, as well as static pressure distributions on the wing and on the fuselage of the wing/body configuration show a satisfying long term repeatability over a time period of 18 years. Flow features at the wing/body junction of the wing/body and the wing/body/fairing configuration have been analyzed in more detail by evaluating inboard pressure distributions as well as pressure sensitive paint to determine the complete pressure field on the wing surface and link it to the information of the pressure distributions. The aerodynamic impact of the side-of-body fairing has been analyzed for design and off design conditions given by a Mach number of 0.80. At off-design conditions adverse drag increments have been observed for the configuration with fairing. Oil flow visualization pictures have been used to investigate the size and formation of the side-ofbody separation. It is shown that the flow inside the separation bubble is mainly fed by the lower wing surface flowfield. For unsteady pressure measurements Kulites sensors have been inserted at the side of the body to analyze the degree of unsteadiness inside and in the vicinity of the separation bubble that is present in the rear part of the wing/body junction.

Computation of Aerodynamic Coefficients for the DLR-F6 Configuration using MEGAFLOW

Navier-Stokes calculations, obtained with software from the MEGAFLOW project, are presented for the Airbus-like DLR-F6 configuration at cruise flight conditions. Results for the wing-fuselage geometry are available for structured grids of different sizes up to 16 million grid cells. The Baldwin-Lomax turbulence model is used and the influence of the numerical dissipation is analyzed. It is demonstrated that even 16 million grid cells are not sufficient to reach a fully grid converged solution for this configuration at transonic flow conditions, if standard non-adapted grids are used. The influence of the numerical dissipation on the pressure distribution is very small for the fine grid. However, the drag coefficient shows a variation of 2.5% for different levels of numerical dissipation. For the configuration with pylon and nacelle, grids with 3.3 and 3.8 million cells are used. In addition to the Baldwin-Lomax model, first computations using the k-ω turbulence model from Wilcox are presented. The pressure distributions as well as the lift and drag coefficients are compared to wind tunnel measurements for both configurations. The comparison shows a good quality of the numerical results.

Comparison of the NASA Common Research Model European Transonic Wind Tunnel Test Data to NASA Test Data

Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility, the NASA Ames 11-ft wind tunnel, and the European Transonic Wind Tunnel. In the NASA Ames 11-ft wind tunnel, data have been obtained at only a chord Reynolds number of 5 million for a wing/body/tail = 0 degree incidence configuration. Data have been obtained at chord Reynolds numbers of 5, 19.8 and 30 million for the same configuration in the National Transonic Facility and in the European Transonic Facility. Force and moment, surface pressure, wing bending and twist, and surface flow visualization data were obtained in all three facilities but only the force and moment, surface pressure and wing bending and twist data are presented herein.