Search

The need of high order boundary discretization in case of a high order code, e.g. a Discontinuous Gaierkin (DG) Code, has already been demonstrated in the literature. Bassi and Rebay extended the DG method to solve the Navier-Stokes equations for laminar and 3D turbulent flow. In the present paper an extension will be provided to include both three-dimensional flows and curved elements to properly represent three-dimensional bodies with curvature. Some results, like Sows around a sphere and around an Onera M6 wing, are calculated and compared with experiments. [ABSTRACT FROM AUTHOR]

Detailed numerical shape optimization will play a strategic role for future aircraft design. It offers the possibility of designing or improving aircraft components with respect to a pre-specified figure of merit. Optimization methods based on exact derivatives of the cost function with respect to the design variables still suffer from the high computational costs if many design variables are used. However, these gradients can be efficiently obtained by the solution of so-called adjoint flow equations. Furthermore, it is possible to derive these adjoint solvers in an automated fashion by the use of so-called automatic differentiation (AD) tools. In the present paper the efficient and automated differentiation of an entire design chain, including the flow solver, is presented. As test case for AD-generated adjoint sensitivity calculations an inviscid RAE2822 airfoil is chosen under transonic flight conditions for drag reduction. [ABSTRACT FROM AUTHOR]

Flow structure in the annular section of the inlet system of a tuboannular swirl combustor with respect to the swirl intensity influence was investigated computationally complementary to the recent experimental study by Palm et al. [1]. In addition to the non-swirling case, two different swirling configurations corresponding to the swirl numbers S = 0.6 and 1.0 were considered. The simulations were performed by using Large Eddy Simulation (LES) method and a two-layer model scheme hybridizing a near-wall k —ε RANS (Reynolds-averaged Navier Stokes) model covering the wall layer and LES method in the outer layer employing Smagorinsky model. Special attention was devoted to the position of the interface. An in-depth analysis of the mean velocity and turbulence fields reveals an increasingly asymmetric axial velocity profile in the annular pipe and an appropriately shaped profile of the Reynolds stress components corresponding to the enhanced turbulence production in the outer part of the concentric annulus. The present study also aimes at generation of reliable swirling inflow data for future LES of the flow in the combustor flue. [ABSTRACT FROM AUTHOR]

To investigate the development of blade tip vortices under different rotor conditions like thrust and rotational speed, both 2C and 3C PIV measurements were performed on a 40% Mach scaled Bol05 model rotor in hover condition. The vortices were traced from the trailing edge of the blade up to half a revolution behind the blade with azimuth steps between 1° to 10° and different spatial resolutions. [ABSTRACT FROM AUTHOR]

In the present paper we investigate the aeroacoustic characteristics of the flow over a forward facing step. The purpose is to identify relevant regions which are responsible for the noise generation. Correlation measurements are carried out for the pressure fluctuations on the wall, velocity characteristics of the flow and the recorded far-field acoustic pressures. The application of the correlation techniques provides a useful tool in the characterisation of the flow-induced sound sources. For further understanding, flow and acoustic experiments are conducted using different Reynolds numbers and radii of curvature. [ABSTRACT FROM AUTHOR]

Sound generation downstream the nozzle end of a subsonic laminar jet has been investigated using two-dimensional direct numerical simulations (DNS). The nozzle end is modeled by a finite flat plate with Mach numbers of MaI = 0.8 above and MaII = 0.2 below the splitter plate. Behind the nozzle end, a combination of a wake and mixing layer develops. Due to the high amplification rates, disturbances saturate before a pure mixing layer occurs. Non-linear generation mechanisms produce higher harmonic disturbances in the upper boundary layer, resulting in an only quasi periodic solution. The main acoustic sources correspond to the positions of vortex pairing. Broadband noise is emitted instead of tonal noise, known from the pure mixing layer without a splitter plate [4]. [ABSTRACT FROM AUTHOR]

In this study the flow field and the acoustic field over a high-lift configuration consisting of a deployed slat and a main wing is numerically analyzed. The flow parameters are Ma = 0.16, Re = 1.4 * 106, αSlat = 23° and αgeo = 13°. The flow data, which are computed via a large-eddy simulation, provide the distributions being plugged in the source terms of, e.g. the acoustic perturbation equations (APE) or the Ffowcs-Williams and Hawkings (FWH) equation, to compute the acoustic near field. The analysis shows the interaction of the shear layer of the slat trailing edge and the slat gap flow to generate higher vorticity than the main airfoil trailing edge shear layer. Thus, the slat gap is the dominant noise region. [ABSTRACT FROM AUTHOR]

Trailing-edge noise data from an experimental study on a flat-plate and on a two-dimensional NACA 0012-like airfoil are presented. Within extensive measurements in the Aeroacoustic Wind Tunnel Braunschweig various flow-permeable (comb-type) edge modifications were tested with respect to their noise reduction capability. Among numerous design parameters a narrow slit width was identified as major requirement for a low-noise trailing-edge design whereas flexibility of the comb material was found to be not imperative for such edge noise reduction means. Because both boundary-layer tripping and trailing-edge thickness show significant influence on trailing-edge noise, a detailed description of these effects is provided to enable an accurate interpretation of corresponding noise test data. [ABSTRACT FROM AUTHOR]

Direct numerical simulations (DNS) are carried out to investigate the effect of effusion cooling by blowing through spanwise slits and discrete holes in a laminar flat-plate boundary layer at various Mach numbers. The comparison with experimental data for a Mach-2.67 boundary layer with a cool wall and a spanwise slit shows good agreement. For an adiabatic Mach-6 boundary layer, the focus of the study is the effectiveness of various effusion-cooling configurations (slits and holes), the generated vortices and shear-layer systems. In terms of cooling effectiveness slits are preferable over (a few) holes. [ABSTRACT FROM AUTHOR]

A large-eddy simulation using preconditioning is described and applied to investigate continuous tundish flow. The Navier-Stokes equations for compressible flow are approximately solved using an implicit dual-time stepping scheme combined with low Mach number preconditioning and a multigrid acceleration technique. The impact of a turbo-stopper on jet spreading, jet impingement on the wall, wall jets, and steel quality is investigated. The numerical analysis shows the turbostopper flow to generate a more homogeneous tundish flow and as such a higher steel quality can be expected. [ABSTRACT FROM AUTHOR]

In the present study, Detached Eddy Simulation is applied to the supersonic axisymmetric base flow. Computations are done on coarse and fine unstructured grids to examine the influence of grid refinement on flow predictions. The final results are compared with available experimental data of Herrin and Dutton, which includes boundary layer velocity profile, base pressure, centerline velocity profile, Mach and turbulent kinetic energy distributions. The influence of grid refinement is observed and the reasons are discussed. [ABSTRACT FROM AUTHOR]

A Seperation of Variables (SoV) scheme has been developed in order to obtain a direct solution to the Poisson problem arising from Chorin's projection method within rectangular enclosures. The algorithm is compared to the previously used FFT with periodic boundaries and methods to improve the performace of the required transformations are presented. Finally, first results of thermal convection obtained in a closed box geometry are compared to results obtained in a periodic rectangular cell. [ABSTRACT FROM AUTHOR]

This paper deals with the effects of heat release and compressibility on temporally evolving, turbulent mixing layers. Direct Numerical Simulations (DNS) of such layers at two different convective Mach numbers are performed with and without combustion which allows to study the effects of heat release and compressibility separately and combined. It is shown that both, compressibility and heat release, dampen the turbulence activity and lead to a reduced growth of the mixing layer. Alterations in pressure fluctuations are a main reason for the changes. The effects of compressibility are not as strong in the reacting mixing layer as in the inert one. A significant difference between the inert and the reacting mixing layers at high convective Mach number is that entropic density fluctuations prevail over the acoustic ones when reaction takes place while both contribute to nearly equal parts in the non-reacting compressible flow. [ABSTRACT FROM AUTHOR]

Simulations and experiments were performed to capture the spatio-temporal flow field around a finite circular cylinder mounted on a ground plate. In order to provide a combined database and testcase for future simulations and experiments, the flow is investigated using state-of-the-art techniques with a high resolution in time and space, namely Large-Eddy Simulation and Detached-Eddy Simulation for the numerics and time-resolved PIV as well as LDA for the measurements. The predicted time-averaged and unsteady flow field from simulations corroborate well the experiments, giving new insights into the complex turbulent separated flow behind a quite simple geometry. [ABSTRACT FROM AUTHOR]

Results of steady and unsteady RANS computations of the flow around the Stratospheric Observatory for Infrared Astronomy SOFIA are presented. The observatory consists of a Boeing 747 SP with an open port in the fuselage to house a 2.5 m infrared telescope for astronomic remote sensing purposes. Results of CFD-simulations, carried out at the University of Stuttgart, show that URANS is able to capture the main effects of the unsteady cavity flow and acoustics inside the SOFIA telescope port. Pressure spectra taken at several points on the telescope's surface point out the presence of unsteady pressure fluctuations at discrete frequencies. Results compare well with experimental data, generated by NASA in wind tunnel investigations with a 7% model of the SOFIA aircraft. [ABSTRACT FROM AUTHOR]

The SSG/LRR-ω differential Reynolds stress model has been implemented into DLR's unstructured flow solver TAU. The model is a blend between the Speziale-Sarkar-Gatski model (SSG) in the far field and the Launder-Reece-Rodi model (LRR) near walls, combined with Menter's baseline ω-equation for the length scale. The implementation has been checked for the transonic flow around the RAE 2822 airfoil, comparing with the results obtained with DLR's structured flow solver FLOWer. In the following the model is applied to the L1T2 three-element airfoil at high incidence, demonstrating the model's applicability to 2D high-lift problems. Furthermore results are shown for the transonic flow around the two wings of the Third AIAA Drag Prediction Workshop and the ONERA M6 wing. [ABSTRACT FROM AUTHOR]

Pod drives are modern outboard ship propulsion systems with a motor encapsulated in a watertight pod. The motor's shaft is connected directly to one or two propellers. The whole unit hangs from the stern of the ship and rotates azimuthally, thus providing thrust and steering. In the past decade pod drive propulsion has developed to be a reliable and serious alternative to conventional ship propulsion, especially for cruiser liners [1,2]. The overall efficiency of the system is determinant for the economical success of the vessel and for this reason it is necessary to improve the efficiency of the propulsor recurring to non-conventional solutions. [ABSTRACT FROM AUTHOR]

The investigations of the aerodynamic performance of new locally swept lattice wings include numerical simulations as well as wind tunnel measurements. The investigations were performed at free-stream Mach numbers from 2 to 6 for angles of attack varied from 0 to 10 degrees. The performance of the lattice wings was assessed on the basis of the zero-lift drag and lift-to-drag ratio. The numerical and experimental results show that the novel design of the lattice wings has distinct advantages in comparison to the conventional unswept configurations. Compared to conventional lattice wings the maximum benefit e.g. of the zero-lift total drag for the investigated locally swept lattice wings is of the order of 30%-40%. [ABSTRACT FROM AUTHOR]

In the present investigation the DLR TAU code is extended to support future experimental investigations of magnetohydrodynamic effects in high temperature hypersonic flows. According to the conditions in the High Enthalpy Shock Tunnel Göttingen (HEG) the first steps in enhancing the TAU code are the implementation of a source term formulation of electromagnetic forces and the calculation of the electrical conductivity of air as a gas mixture in chemical non equilibrium. To verify the source term implementation a perfect gas study related to numerical simulations from Poggie and Gaitonde is conducted and shows reasonable agreement. Applied to the experimental conditions the model predicts a noticeable increase of the shock stand off distance. [ABSTRACT FROM AUTHOR]

The total pressure losses in a compressor stage are strongly affected by sec-ondary flow effects such as corner separation. Experiments in a highly loaded com-pressor cascade were performed to minimize the losses using a grooved surface on the side walls under cross flow conditions. This geometry hampers the secondary flow in the cascade in such a way, that the corner separation is less pronounced. The total pressure losses are reduced by 9.8% and the flow turning angle is increased. The experiments were accomplished at Reynolds numbers up to Re = 0 56 · 10 (based on 40 mm chord of the vane profile) and Mach numbers up to M = 0.67. To understand the flow physics of the crosswise grooves, additional experiments at lower Reynolds numbers were performed. [ABSTRACT FROM AUTHOR]

The aerodynamic loads for a missile radome cruising near ground level are investigated. The investigation consists of experiments at Mach 6 performed in the High Enthalpy Shock Tunnel Göttingen (HEG) and CFD calculations using the TAU code of DLR. The first approach duplicated an hypersonic flight at 11 km altitude for some milliseconds with the appropriate Mach number, Re number, flow velocity, pressure and heat load. The experimental results for the heat transfer on the radome fit to the CFD data including the transition from laminar to turbulent. [ABSTRACT FROM AUTHOR]

A blunt-nosed supersonic projectile, equipped with a cavity, is a good alternative to reduce high nose heating rates preventing surface melting followed by ablation. For this purpose, tests were done in ISL's shock tunnel STB. This facility allows reproducing atmospheric flow conditions present during a missile flight at various altitudes. The flow was visualized by shadowgraphs and the heat flux densities were measured with thermocouples at the bottom of various cavity geometries. A numerical simulation was also carried out using the FLUENT code. The comparison between numerical results and measurements is quite satisfactory. An important result of this study is that the deepest cavity has the smallest heat flux. [ABSTRACT FROM AUTHOR]

The given paper summarises the status of the post-flight analysis of the SHarp Edge Flight EXperiment (SHEFEX). It focuses on the comparison of numerical and experimental surface pressures as well as the assessment of the free stream vector applying the flight mechanic description of the Digital Miniature Attitude Reference System (DMARS) platform. The results point out that the extraction of angles of attack and sideslip is much more demanding than expected. [ABSTRACT FROM AUTHOR]

This paper outlines the theory of radical farming in scramjets and describes the experimental scramjet model that was designed to investigate it. Experiments were conducted at two conditions; a 3MJ/kg condition corresponding to Mach 7.9 flight at an altitude of 24km and a 4MJ/kg condition corresponding to Mach 9.1 flight at an altitude of 32km. The results are presented as pressure distributions on the flowpath wall and specific impulse estimates. [ABSTRACT FROM AUTHOR]

During the past years, the Centre National d'Etudes Spatiales (CNES) has initiated several studies concerning the first generation experimental vehicle Pre-X, which is dedicated to mastering glided atmospheric re-entry. CNES has committed at the beginning of 2005 to an important phase with the objective of reaching a preliminary design review by mid 2006. In the framework of phases A2 and B of the Pre-X program the German Aerospace Center (DLR) carried out wind tunnel experiments on a scaled Pre-X vehicle in the High Enthalpy Shock Tunnel Gottingen (HEG) as well as numerical investigations using the DLR TAU code. For a 20° body flap deflection, transitional reattachment was experimentally observed. Additionally, the experiments and the CFD indicated a heating region at the end of the flaps that originated from the impingement of the interference between the bow shock and the shock created by the re-attachment. [ABSTRACT FROM AUTHOR]

One of the major goals of flow control is the reduction of energy consumption in wall bounded flows by minimizing the viscous drag. In the present work it is shown that the turbulent dissipation needs to be minimized in order to obtain energy savings. Analytical considerations lead to the conclusion that this can be achieved by forcing near-wall turbulence to an axisymmetric state. To confirm this finding direct numerical simulations were carried out in which the boundary conditions were such that near-wall fluctuations were forced towards axisymmetry. Additionally, a surface structure that minimizes turbulent dissipation at the wall was designed and tested experimentally. Both, numerical and experimental investigations yield significant drag reduction. [ABSTRACT FROM AUTHOR]

This paper describes the development of an active control system for delaying laminar-turbulent transition in compressible flows. Convective boundary layer instabilities are detected during their linear amplification stage by a reference sensor and reduced downstream through destructive interference of artificial counterwaves. [ABSTRACT FROM AUTHOR]

Flight tests were performed in order to learn more about free-stream excited Tollmien-Schlichting waves and the role they play in boundary layer transition. Therefore, a multielement hot-film array, placed on the right hand wing of the flying testbed LFU-205, was used to sense the skin friction fluctuations provoked by this type of instability while propagating in the laminar boundary layer. The experimental data acquired were subsequently analyzed with special emphasis on the evaluation of the initial amplitudes of Tollmien-Schlichting waves. For the streamwise disturbance velocity component their magnitude was determined to be in the order of 10-9 times the boundary layer edge velocity. [ABSTRACT FROM AUTHOR]

Results of an experimental study of controlled laminar-turbulent transition in a smooth axisymmetric diffuser at 7800 < ReD1 < 10600 based on the inlet diameter D1 and the bulk velocity $$ u_m = \tfrac{2} {{D_1 }}\int_0^{D_1 /2} {u(r)dr} $$ are presented. The inlet flow is an incompletely developed laminar pipe flow with a clear boundary-layer shape (boundary layer thickness δ99/D1 ≤ 0.3). The smooth diffuser contour causes the formation of a closed pressure-induced laminar separation bubble, which is notedly receptive for small-scale disturbances. Upstream of the diverging geometry, controlled periodic perturbations are introduced into the boundary layer. The instability of the local velocity profiles with an inflection point causes a massive growth of instability waves within the shear layer leading to a turbulent breakdown of the laminar flow. For two different perturbation modes, the resulting mean flow field as well as the velocity fluctuations are measured by means of a Laser Doppler Velocimetry (LDV) and Particle Image Velocimetry (PIV) system in detail. [ABSTRACT FROM AUTHOR]

Direct numerical simulation of a pressure-induced short laminar separation bubble developing on a flat plate has been carried out. Transition in this bubble was triggered by small disturbance input with a fixed frequency and fixed spanwise wave number. The resulting short bubble was shown to be converged in time to a statistically steady state, while possessing essential features of short laminar separation bubbles as reported in the literature. In the present case disturbance input is required to maintain a short bubble. Switching off this disturbance input yields a growing separation bubble. This phenomenon is denoted as bubble bursting, since indication is found that the bubble develops towards a long-bubble state. [ABSTRACT FROM AUTHOR]

The method of temperature-sensitive paint (TSP) adapted to cryogenic wind tunnels, i.e. cryoTSP is used to detect the natural laminar-turbulent transition of boundary layers in high speed flows. Besides pure transition detection, cryoTSP images provide information on the transition zone, the characteristics of transition, and on the type of instability. A method is proposed to determine the transition zone out of cryoTSP result images for 2D-wings, where the transition process is caused by Tollmien-Schlichting instabilities. [ABSTRACT FROM AUTHOR]

This study shows the application of the criterion of Pfenninger for the prediction of the influence of geometric variations at the slat of a multi-element airfoil on the occurrence of attachment line transition (ALT). Additionally the use of criteria for the prediction of relaminarization based on the results of 2D RANS computations is examined. Flow calculations are performed for a representative wing section of a three-dimensional swept and tapered high-lift wing for which the occurrence of ALT is assumed at higher Reynolds numbers. Systematic geometry variations are performed to investigate the influence on the predicted onset of ALT and the influence on aerodynamic characteristics in order to identify possibilities for the design of a slat with delayed ALT onset. From this the most sensitive parameters for ALT transition and aerodynamic performance have been identified. [ABSTRACT FROM AUTHOR]

A Reynolds-averaged Navier-Stokes solver, a laminar boundary-layer code and different transition prediction methods for the prediction of Tollmien-Schlichting and cross flow instabilities were coupled for the automatic prediction of laminar-turbulent transition on general 3-dimensional aircraft configurations during the ongoing flow computation. The procedure is applied to a two-dimensional three-element high-lift airfoil configuration which is characterized by the existence of laminar separation bubbles using different operation modes of the procedure. [ABSTRACT FROM AUTHOR]

Recently, Runge-Kutta/Implicit methods were proposed for the solution of the Euler and Navier-Stokes equations, allowing a reduction of computation times by about half an order of magnitude compared to methods presently in use. In this contribution the efficiency of such methods is further enhanced, first by reducing the number of Runge-Kutta stages, and second by introducing direct line-solves in the direction of maximum stiffness. Compared to a well tuned, standard reference code, for high-Reynolds number flows computation times are reduced by more than an order of magnitude. [ABSTRACT FROM AUTHOR]

The solution of the discrete adjoint equations for an unstructured finite volume compressible Navier-Stokes solver is discussed. In previous work fixedpoint iterations taken from the non-linear method - suitably adjointed - were applied to the adjoint problem. Here it is seen that there are often situations in which these iterations can not be expected to converge. To address this the Recursive Projection Method is developed as a stabilizer, and then used to perform an eigenmode analysis of attached and separated flow on a single geometry, allowing identification of flow regions that were unstable under the basic iteration. Finally an adjoint based optimization with 96 design variables is performed on a wing-body configuration. The initial flow has large regions of separation, which are significantly diminished in the optimized configuration. [ABSTRACT FROM AUTHOR]

Vortex dominated flows appear in many flow simulations such as wake turbulence of an aircraft or a delta wing at a high angle of attack. For detailed investigations of vortex breakdown, vortex interactions or tracing vortex cores, an automated grid adaptation with suitable vortex indicators is essential. Physical indicators, e.g. the vorticity magnitude or the total pressure loss, are in most cases not sufficient for correctly identifying a vortex core. [ABSTRACT FROM AUTHOR]

A one rotational degree of freedom and a six degree of freedom flight mechanics module are coupled to DLR's RANS solver FLOWer. The algorithms used to solve the equations of motion are described. The 1-DOF module is validated for a freely rolling delta wing. A very good agreement with experimental data is observed. The capability of the 6-DOF module is demonstrated by a simulation of a store separating from a wing. The computation uses the overset grid technique to handle the movement of the bodies in relative motion. [ABSTRACT FROM AUTHOR]

The use of unsteady adaptation with the DLR-TAU Navier-Stokes solver is presented as a method of improving the modelling of flows where the aerodynamic performance of a body is determined by the action of moving localised regions of high-gradient flow. Examples are presented of transonic limit cycle oscillation and dynamic stall. First results indicate that good grid convergence can be achieved without necessarily requiring that the flow around the specific airfoil is well understood in advance. [ABSTRACT FROM AUTHOR]

Mesh deformation for unstructured grids is usually solved employing algorithms based on a spring analogy or on a linear elasticity analogy. These methods require solving discrete equations for the given grid using iterative methods. The computational effort for these methods is not negligible compared to the effort required by the flow solver. In order to optimise the efficiency of the computational chain for coupled simulations of fluid structure interaction an alternative algebraic method was developed for the DLR TAU-Code. The TAU-Code is parallelized by domain decomposition using the message passing concept based on MPI such that each process works on one partition of the grid only. In order to employ the grid deformation on the grid partitions distributed over the processes, a parallel method is required. This paper describes the parallel grid deformation method and shows its efficiency and robustness for the DLR F6 geometry. Further applications are shown in order to demonstrate the grid deformation for dynamic aeroelastic simulations. [ABSTRACT FROM AUTHOR]

In this paper the recently developed space-time expansion discontinuous Galerkin (STE-DG) approach for the two dimensional unsteady compressible Navier-Stokes equations is presented. The basis of the scheme is a weak formulation of the Navier-Stokes equations, where special care of the second order terms is taken. The spatial polynomial of the DG approach is expanded in time using the so called Cauchy-Kovalevskaya (CK) procedure. With a polynomial of order N in space the CK procedure generates an approximation of order N in time as well yielding a scheme of order N+1 in space and time. The locality and the space-time nature of the presented method give the interesting feature that the time steps may be different in each grid cell. Hence, we drop the common global time steps and propose for a time-accurate solution that any grid cell runs with its own time step determined by the local stability restriction. In spite of the local time steps the scheme is conservative, fully explicit, and as in the DG approach the polynomial order could be chosen arbitrarily, the scheme is theoretically of arbitrary order of accuracy in space and time for transient calculations. [ABSTRACT FROM AUTHOR]

Starting from a mixed fidelity multiple discipline analysis process for high-speed transport aircraft, a multiple discipline optimisation scenario is set-up aiming to obtain minimum weight configurations. A new, more realistic objective function driving the optimisation is implemented. A realistic flight mission for a supersonic transport aircraft (from take-off to landing) is modelled using that process. For supersonic cruise conditions -were most flight time is spent-the disciplines aerodynamics and structures are modelled at high fidelity while other contributing disciplines and flight phases are modelled less accurate. For a generic long-range mission the impact of the supersonic cruise-speed on the minimal take-off weight (enabling to cruise this long range mission) is addressed. [ABSTRACT FROM AUTHOR]

An investigation of both an internal and an external stress wave force balance that measures forces and moments on a re-entry type configuration was undertaken. Simulations performed using a finite element code suggest superior recovery of the applied forces and moments with an external force balance configuration. This is hypothesized to be due to the increased balance stiffness and decoupling between force components that can be achieved with an external type configuration. The influence of an external force balance as compared to an internal force balance on the model loads was determined to be negligible. [ABSTRACT FROM AUTHOR]

The flow over a finite-height cylinder of aspect ratio H:D=2:1 has been studied by means of oil-film visualization, pressure and heat flux fluctuation measurements. The measurements were concentrated on the diameter-based Reynolds number of Red = 2 x 105. A novel heat flux gauge with a high spectral resolution was used to investigate the heat flux fluctuation on the surface of a slightly heated cylinder. Comparative investigations showed qualitative correlations between the pressure and heat flux fluctuations around the cylinder and on its free-end. Several flow features could be identified by means of time-resolved surface measurements. [ABSTRACT FROM AUTHOR]

The Background Oriented Schlieren Method (BOS) and the Laser Light Sheet technique (LLS) were successfully adapted to the cryogenic European Transonic Windtunnel (ETW) and applied to wing tip vortex investigations. Applying optical measurement techniques to ETW is difficult, because of the extreme wind tunnel conditions (temperature, pressure, poor optical access). [ABSTRACT FROM AUTHOR]

In this paper a new approach of CFD supported wind tunnel testing is presented based on investigations of the DLR project ForMEx [1]-[3]. The numerical simulation and respectively the analysis of the wind tunnel experiment considering all geometrical and aerodynamic conditions show improvements of today's wind tunnel testing techniques which is outlined in this paper for the wind tunnel DNW-NWB. [ABSTRACT FROM AUTHOR]

The Pressure Sensitive Copolymer Coating (PSC) with piezoelectric properties is introduced as a new surface measurement technique. In contrast to conventional surface measurement techniques the PSC combines high spatial and temporal resolution, and can be applied to arbitrarily formed surfaces. The linear relationship between the mechanical pressure and the electrical voltage due to the piezoelectric effect is established by calibration measurements. Furthermore, first applications in transition experiments in incompressible as well as in compressible flows are presented. Finally, investigations of the flow around a wall mounted cylinder are presented. [ABSTRACT FROM AUTHOR]

A flat plate boundary layer experiment is designed to examine the physical mechanism and the performance of vortex generator jets at Reynolds numbers up to Reθ = 20000. The investigation of vortex generator jets at this flow state is of fundamental importance for the assessment of flow control at takeoff and landing conditions. First results show that the optimal skew angle is about β = 15° for the slot actuator. The formation of the longitudinal vortex structures can be described in detail by the derived vortex topology based on the measurement results. [ABSTRACT FROM AUTHOR]

Shock control bumps are a means to reduce wave drag that occurs at the upper limit of civil transport aircrafts flight envelope. An SCB was optimized for and applied to the rigid wing-body DLR F11 model. The effect of the SCB on the trim drag of the F11 configuration with an attached horizontal tail plane was investigated. Flexibility of the wing-body model was considered and the influence of aeroelasticity on the SCB performance was examined. RANS simulations with the DLR FLOWer code showed, that both the influence of trimming as well as of aeroelasticity is negligible for SCB design. [ABSTRACT FROM AUTHOR]

Numerical 2D simulations with a RANS flow solver are conducted to find the aerodynamic sensitivities of a gapless high-lift system. The investigated high-lift configuration is an airfoil which utilises trailing edge blowing. A small fraction of the engine flow is used for circulation control. The air is blown from a slot directly upstream of the flap and thus the flow over the flap can bear large adverse pressure gradients without separation. It was found that the use of circulation control yields lift coefficients which are comparable or superior to those generated by conventional high-lift systems. The promising results of the 2D simulations motivate applications to a wing-body configuration. The results show that a gapless high-lift system equipped with circulation control has the ability to provide sufficient lift for take off, climb and landing. [ABSTRACT FROM AUTHOR]

In this work retrofit winglets are designed for a transonic aircraft. The used geometry is a generic twin engine aircraft. In a first step, winglets are designed using a lifting line method along with RANS solutions. L/D is optimized by taking into account a certain wing root bending moment reserve for the reference wing. The final analysis includes wing deformation studies by means of fluid-structure coupling. Therefore a finite element model has been developed with respect to standard loadings of certification authority. Using the fluid-structure coupling process RANS solutions with deformed wing shapes are obtained for cruise conditions in order to determine the influence of deformation on performance and for a 2.5g load case in order to evaluate ultimate structural loading. Comparing the results of the rigid wing with the deformed wing the wing root bending moment and the bending moment of the device is clearly reduced for the deformed wing. Thereby the advantage to which an implementation of the described method for future design processes would lead becomes apparent. [ABSTRACT FROM AUTHOR]