DLR TAU Code Research Articles

Design of Input Signal for System Identification of a Generic Fighter Configuration

This article investigates the design of time-accurate input signals in the angle-of-attack and pitch rate space to identify the aerodynamic characteristics of a generic triple-delta wing configuration at subsonic speeds. Regression models were created from the time history of signal simulations in DoD HPCMP CREATETM-AV/Kestrel software. The input signals included chirp, Schroeder, pseudorandom binary sequence (PRBS), random, and sinusoidal signals. Although similar in structure, the coefficients of these regression models were estimated based on the specific input signals. The signals covered a wide range of angle-of-attack and pitch rate space, resulting in varying regression coefficients for each signal. After creating and validating the models, they were used to predict static aerodynamic data at a wide range of angles of attack but with zero pitch rate. Next, slope coefficients and dynamic derivatives in the pitch direction were estimated from each signal. These predictions were compared with each other as well as with the ONERA wind tunnel data and some CFD calculations from the DLR TAU code provided by the NATO Science and Technology Organization research task group AVT-351. Subsequently, the models were used to predict different pitch oscillations at various mean angles of attack with given amplitudes and frequencies. Again, the model predictions were compared with wind tunnel data. Final predictions involved responses to new signals from different models. A feed-forward neural network was then used to model pressure coefficients on the upper surface of the vehicle at different spanwise sections for each signal and the validated models were used to predict pressure data at different angles of attack. Overall, the models predict similar integrated forces and moments, with the main discrepancies appearing at higher angles of attack. All models failed to predict the stall behavior observed in the measurements and CFD data. Regarding the pressure data, the PRBS signal provided the best accuracy among all the models.

Plume and wall temperature impact on the subsonic aft-body flow of a generic space launcher geometry

Experimental and numerical simulation of launcher base flows are crucial for future launcher design. In experiments, exhaust plume simulation is often limited to cold or slightly heated gases. In numerical simulations, multi-species reactive flow is often neglected due to limited resources. The impact of these simplifications on the relevant flow features, compared to real flight scenarios, needs to be characterized in order to enhance the design process. Experimental and numerical investigations were carried out within the framework of the SFB/TRR 40 Collaborative Research Centre to study the impact of plume and wall temperature on the base flow of a generic small-scale launcher configuration. Wind tunnel tests were performed in the Hot Plume Testing Facility (HPTF) at DLR, Cologne, using subsonic ambient flow and pressurized air or hydrogen–oxygen combustion as exhaust gases. The tests were numerically rebuilt using the DLR TAU code employing a scale-resolved IDDES approach, including thermal coupling and detailed chemistry. The paper combines the experimental and numerical findings from the SFB/TRR 40 base flow studies and highlights the most prominent influences on the mean flow field, the pressure field, the dynamic wake flow motion, and the resulting aerodynamic forces on the nozzle. High-frequency pressure measurements, high-speed schlieren recordings, and time-resolved CFD results are evaluated using spectral and modal analysis of the one- and two-dimensional flow field data.

Shock–shock interaction on the Reusability Flight Experiment (ReFEx)

AbstractThe Reusability Flight Experiment (ReFEx) is currently under development at the German Aerospace Center (DLR). A coupled experimental and numerical campaign was carried out to investigate the surface heating on the payload geometry during return consisting of a forebody and canard. In this way, numerical tools for a post-flight analysis can be preemptively improved where required. Experiments were undertaken at the High Enthalpy Shock Tunnel Göttingen (HEG) on a 1:4 scale model with the use of temperature sensitive paint on the payload geometry to obtain surface heat flux. The model configuration was varied in angle of attack and canard deflection. Laminar and turbulence-model solvers in the DLR-TAU code were used for the numerical simulations. This investigation focussed on the shock–shock interaction of the nose bow shock with the leading-edge shock of the canards showing significant surface heat flux along the canard. Larger surface heat fluxes were measured in the experiments downstream of the shock interaction on the canard, than obtained from the laminar CFD calculations. This was attributed to transition of the boundary layer within the interaction regions, in the presence of significant adverse pressure gradients. Other flow features along the forebody in the vicinity of the canard were qualitatively matched better by the fully-turbulent numerical solutions than the laminar ones. This work aims to demonstrate the extent to which the numerical and experimental tools assist useful insights into flow phenomena during the return stages of the ReFEx payload geometry, and the aspects for which improvements are required.

Influence of structural elasticity on the aerodynamic characteristics of the common research model with deflected control surfaces

Numerical aerodynamic and aeroelastic investigations are conducted for the Common Research Model (CRM) configuration with deployable aileron control surfaces and a trimmable horizontal stabilizer at cruise and off-cruise conditions. The numerical methodology which is implemented in the multi-disciplinary simulation environment SimServer is based on CFD and coupled CFD-CSM simulations. The DLR Tau Code is applied in order to solve the Reynolds-Averaged-Navier-Stokes (RANS) equations. The Spalart-Allmaras turbulence model with Edwards modification is used for the closure modeling of the three-dimensional RANS equations. A linear modal solver is utilized to calculate structural displacements and rotations of control surfaces are modeled by a Chimera approach. A grid independence study is conducted by using the rigid aerodynamic model of the CRM. The main focus of the analysis lies on the evaluation of the stationary CFD-CSM simulations addressing the influence of structural deformation on aileron effectiveness. Influences of various flow conditions and different fuel tank filling levels on the resulting local and global aerodynamic properties of the CRM are investigated. The reduction of aileron effectiveness as well as the difference in horizontal trim deflection angle if structural elastic effects are considered is investigated for the considered Mach number range and aircraft mass configurations.

Numerical Investigation of an Unmanned Aerial Vehicle Launch from Military Transport Aircrafts

Launching a fixed-wing unmanned aerial vehicle (UAV) out of the cargo hold of a flying transport aircraft (TA) is numerically investigated. A numerical tool chain is established to capture the dynamic motion of the UAV during launch, in which the DLR-TAU Code flow solver is coupled with a flight mechanics tool. A parameter study investigates how the UAV launch speed, position, and orientation relative to the loading platform of two different TAs affect its trajectory and aerodynamic behaviors. Furthermore, the influence of the angle of attack of the TA on the UAV trajectory and pitch angle is analyzed. This information will help design a launch mechanism that ensures a safe separation and aerodynamically stable flight after launch.

Galilean-Invariant Stability-Based Transition Transport Modeling Framework

A new stability-based transition transport modeling framework for unstructured computational fluid dynamics is introduced. Based on a four-equation model published recently by the authors of this paper, new model formulations are derived being Galilean invariant and fully local. Moreover, the complexity of the four-equation model is reduced, yielding a one-equation model that is based on a local pressure-gradient parameter and a three-equation model that is based on an averaged pressure-gradient parameter considering convection (history) effects. Since the model versions are identical except the pressure-gradient parameter, the implementation effort is reduced. To be able to compute test cases in adverse pressure gradient flow regions at low Reynolds numbers, a method was developed to accelerate the turbulence production (downstream of the point of transition onset) that can also be used inside laminar separation bubbles. The models were implemented into the DLR TAU code, the unstructured compressible finite-volume flow solver of the German Aerospace Center (DLR) for external flows. For verification and validation purposes different test cases are shown capturing a wide range of parameters and flow conditions. This includes a test case verifying that the model yields identical results with and without uniform motion (Galilean invariance). Additionally, industrially relevant three-dimensional test cases were computed. This includes a helicopter rotor in vertical flight.

Capsule aerodynamics and shock-wave boundary layer interaction (SBLI) in supersonic and transonic flow

Static force and moment measurements are performed on the MarcoPolo-R aero shape in the trisonic wind tunnel TMK. The static stability behaviour of the capsule is characterized in the Mach number range 0.5le Mle 3.5 reproducing the Mach and Reynolds conditions of the flight trajectory in the supersonic regime. An aerodynamic database is built based on the experimental results. The flow structure around the capsule is visualised in supersonic tests by means of schlieren imaging. Under certain conditions, development of a complex shock system on the leeward side of the inclined capsule is observed. Oil film technique is used to visualise boundary layer phenomena in connection to this shock system. Numerical simulations with the DLR TAU code are performed to support the interpretation of the flow phenomena under these conditions.Graphical abstractSchlieren imaging visualisation of density gradients in supersonic flow. Depending on the test conditions and angle of incidence, a complex shock system is observed on the leeward side of the inclined capsule. Analysis of the shock structure and its impact on aerodynamic coefficients is one subject of the present investigation.

Investigation of transonic aerodynamics on a triple-delta wing in side slip conditions

The vortex-dominated flow around the triple-delta wing ADS-NA2-W1 aircraft is investigated in order to achieve a better understanding of the flow physics phenomena that occur over the aircraft particularly at the transonic speed condition. Both URANS and scale-resolving DDES have been employed in order to explore the range of suitability of current CFD methods. The Spalart–Allmaras One-Equation Model with corrections for negative turbulent viscosity and Rotation/Curvature (SA-negRC) is employed to close the RANS equations, whereas the SAneg-based DDES model is applied in the scale-resolving computations. The DLR TAU-Code is used to perform the numerical simulations. The deficiencies of the URANS results are illustrated and promising improvements are reached employing the SAneg-DDES numerical method. The hybrid method results show great advancement in the prediction of the multiple-delta wing flow by revealing physical aspects which have not been seen from URANS with sufficient accuracy like vortex–vortex interaction and shock-vortex interaction. These phenomena furthermore explain in a clear way the improved prediction of the surface pressure coefficient over the aircraft and consequently of the aerodynamic force and moment coefficients.

Aeroelastic simulations of a delta wing with a Chimera approach for deflected control surfaces

A numerical tool for the computation of aircraft control surface aerodynamics with flexibility effects is presented. The solution is based on coupled Computational Fluid Dynamics (CFD) and Computational Structural Mechanics (CSM) simulations embedded in the multidisciplinary simulation environment SimServer. In SimServer, the DLR-TAU Code is utilized to obtain the CFD solution by solving the Reynolds-Averaged Navier–Stokes (RANS) equations. Structural displacements are computed with a modal solver. The Chimera implementation of SimServer, suited for hybrid grids, is applied to model the control surfaces. Numerical simulations with the flexible Chimera method are performed for the Model53 wing configuration, which is a generic delta wing with a deployed slat as well as an inboard and outboard trailing edge flap. Aerodynamic and aeroelastic simulations at high dynamic pressure q=45 kPa and transonic speed {text {Ma}} = 0.8 are performed for several angles of attack 10^circ le alpha le 25^circ and flap deflection angles -30^circ le delta le 30^circ. The effect of structural deformations on the flow field and control surface effectiveness are analyzed and compared to computations of components treated fully rigid. At the targeted freestream condition M=0.8 and {text {Re}}=15.1 times 10^7, the flow field around the Model53 configuration is characterized by the interaction of vortices and shock waves. The results of the lift and pitching moment coefficient for the rigid and flexible configuration revealed the importance of taking the structural flexibility into account in order to obtain more accurate results for the considered range of flap deflections. Furthermore, the computational effort of the aerodynamic and aeroelastic simulations are evaluated. The increase in computational effort is shown to be adequate for the given increase in accuracy.

An implicit high-order k-exact finite-volume approach on vertex-centered unstructured grids for incompressible flows

We present a k-exact reconstruction method, which can be incorporated into vertex-centered unstructured finite-volume flow solvers to maintain a high-order accurate solution in space. The scheme is combined with a fractional step strategy for the solution of the incompressible Navier-Stokes equations with a fully implicit discretization of a Poisson equation for the pressure correction. This also involves a novel approach for the flux discretization in the k-exact framework. It is shown, that a third order of accuracy can be maintained for the discretization of convective fluxes and a second order of accuracy for diffusive fluxes, even on highly distorted grids. The scheme is implemented into ThetaCOM, a turbulent heat release extension of DLR's TAU code in its combustion version, showing its applicability to be used in a full production flow solver. The improved properties of our approach in terms of performance and accuracy are demonstrated with several benchmarks cases in both two and three spatial dimensions.

Numerical and experimental studies of the hypersonic flow around a cube at incidence

In order to improve predictions of the on-ground casualty risk associated with the uncontrolled atmospheric re-entry of satellites from Low Earth Orbit, there is significant research interest in the development of engineering models of hypersonic heating rates to faceted shapes. A key part of developing such models is generating accurate datasets of the heat fluxes experienced by faceted shapes at various orientations in hypersonic flows. In this work, we use wind tunnel experiments and CFD simulations to study the hypersonic flow around a cuboid geometry at 5° incidence in a Mach 5 flow at Reynolds numbers of 79.5×103, 109×103 and 148×103. The wind tunnel data are obtained in the University of Manchester's High SuperSonic Tunnel and consist of schlieren images and temperature histories collected using infrared thermography. These temperature histories are used to calculate experimental heat fluxes by solving a three-dimensional inverse heat conduction problem. CFD simulations around the same geometry at equivalent free-stream conditions are calculated with the DLR-TAU code. The experimental and CFD results show good agreement both in terms of heat fluxes as well as flow structure. Notable flow structures include wedge-shaped regions of high heat flux which emanate from the windward corners of the cube. Analysis of numerical Q-criterion contours show that these high heat flux regions are caused by vortex structures generated by the expansion at the cube corner. Analysis of the numerical skin friction coefficient shows that even at incidence there is no breakaway separation from the expansion edges of the cube and the flow remains attached throughout. We show that although there is little change in the average heat flux experienced by a cube at 5° incidence to the free-stream compared to one at 0° incidence, there are significant changes in the heat flux contours over the cubes at these two incidences. Finally, we calculate a number of heating shape factors which can easily be implemented in satellite re-entry and demise prediction analysis tools.

Preliminary aerodynamic design of a reusable booster flight experiment

The Reusable Flight Experiment (ReFEx) is an experimental vehicle currently in development by the German Aerospace Center (DLR), which simulates the reentry of a winged reusable booster stage. The topic covered in this paper is the aerodynamic design through the Mach number range from 5.0 to 0.8, including the transonic regime, by CFD simulations using the DLR TAU-Code. For steering and lift generation, the vehicle is equipped with multiple aerodynamic surfaces: canards, wings and a vertical fin. A VSB30 was chosen as the carrier rocket, leading to numerous limiting conditions that had to be fulfilled. The resulting tradeoffs are discussed. Also the process to achieve a configuration that meets all needs regarding stability during ascent and descent, controllability and energy management to autonomously follow a desired trajectory is explained. Since this vehicle is under development, the presented geometry is not necessarily the final shape. However, the completed design iterations strongly indicate that only details will differ between the presented state to the final and flyable ReFEx, scheduled to fly in 2022.

Experimental and Numerical Analysis of a Streamwise Vortex Downstream of a Delta Wing

The present paper analyzes the vortical flow downstream of a delta wing. The delta wing is designed to generate a vortex that resembles the vortex of a nacelle strake. The experimental results of stereoscopic particle image velocimetry measurements performed at different positions downstream of the delta wing display the streamwise development of the vortex. Details of the postprocessing of the measurement data and a quantification of the uncertainties are presented. The evaluation of the mean axial and tangential velocity components and Reynolds stresses provides insights to the vortex characteristics and their development in the streamwise direction. In addition, results of numerical simulations with the DLR TAU-code applying the Menter-Shear Stress Transport (SST) eddy viscosity model and the Speziale-Sarkar-Gatski/Launder-Reece-Rodi (SSG/LRR)- Reynolds stress model show the ability of the turbulence models to capture the vortex development. The experimental results indicate a preservation of the vortex strength and structure downstream of the delta wing. The Menter-SST model does not predict this preservation; rather it computes a rapid decay of vortex strength and an increase in vortex size. In contrast, the results of the SSG/LRR- model are in good agreement with the experiments concerning the mean flow development. However, both turbulence models underpredict the turbulence in the vortex core.

Responses of vertical axis wind turbines to gusty winds

Gust is a common atmospheric condition encountered by wind turbines. Despite the presence of a vast amount of literature on this topic, few of them involve lateral gust influence. Due to this motivation, this paper presents an investigation on the effects of sinusoidal gusts on a three-bladed vertical axis wind turbine under varying gust parameters including gust direction, velocity amplitude, and frequency. The chimera mesh technique is used to model the real rotation of the rotor, as well as the resolved gust approach model in the DLR (German Aerospace Center) TAU code for gust simulation. Both the general aerodynamic computational fluid dynamics model and the gust model are validated before the following simulations. Numerous new flow phenomena and physics are revealed. The influences of gust on the rotor power output and flowfield characteristics are discussed and analyzed in detail. The findings in this study may be helpful for some practical wind engineering applications, such as atmospheric influence evaluation and field site selection.

DLR TAU-Code uRANS Turbofan Modeling for Aircraft Aerodynamics Investigations

In the context of an increased focus on fuel efficiency and environmental impact, turbofan engine developments continue towards larger bypass ratio engine designs, with Ultra-High Bypass Ratio (UHBR) engines becoming a likely power plant option for future commercial transport aircraft. These engines promise low specific fuel consumption at the engine level, but the resulting size of the nacelle poses challenges in terms of the installation on the airframe. Thus, their integration on an aircraft requires careful consideration of complex engine–airframe interactions impacting performance, aeroelastics and aeroacoustics on both the airframe and the engine sides. As a partner in the EU funded Clean Sky 2 project ASPIRE, the DLR Institute of Aerodynamics and Flow Technology is contributing to an investigation of numerical analysis approaches, which draws on a generic representative UHBR engine configuration specifically designed in the frame of the project. In the present paper, project results are discussed, which aimed at demonstrating the suitability and accuracy of an unsteady RANS-based engine modeling approach in the context of external aerodynamics focused CFD simulations with the DLR TAU-Code. For this high-fidelity approach with a geometrically fully modeled fan stage, an in-depth study on spatial and temporal resolution requirements was performed, and the results were compared with simpler methods using classical engine boundary conditions. The primary aim is to identify the capabilities and shortcomings of these modeling approaches, and to develop a best-practice for the uRANS simulations as well as determine the best application scenarios.

DLR TAU Simulations for the Second AIAA Sonic Boom Prediction Workshop

The accurate prediction of the sonic boom is essential for the development of a future low-boom supersonic aircraft. In this paper, geometries and grids provided by the Second AIAA Sonic Boom Prediction Workshop are used to improve a best practice for predicting near-field pressure signatures with the DLR TAU code. The workshop-provided low-boom geometries with different complexities, such as an axisymmetric body, a delta-wing body, and a full aircraft configuration with a cruciform tail and an engine. Results for the DLR TAU simulations on the workshop-provided grids are presented for different grid refinement levels using both Euler and Reynolds-averaged Navier–Stokes simulations. In addition to the provided grids, mixed-element grids with varying setups are generated. The positions of the shocks and expansions are essentially identical for all medium to fine grids but vary for very coarse grids due to the surface triangulation. The pressure magnitude at shocks and expansions is higher for simulations on fine grids compared to coarse grids. It is shown that the influence of the viscosity on the pressure signatures is more significant than the influence of the grid resolution. The agreement between TAU solutions and other participants’ solutions is good.

Enhanced Maneuverability of a Delta-Canard Combat Aircraft by Vortex Flow Control

Delta-canard shapes allow for maneuverable, supersonic combat aircraft. A configuration of medium aspect ratio was investigated to extend its maneuverability to very high angle of attack. Here, leading-edge vortex flows with intense interaction and breakdown effects dominate the flow. To appreciably enhance the maneuverability beyond limits, aerodynamic means had to be found. In-depth understanding of the interactions of the vortex systems was necessary to improve the aircraft capabilities. Unsteady Reynolds-averaged Navier-Stokes simulations simulations were performed via the unstructured DLR-Tau code. The aircraft was modeled without compromising geometrical details. It was found that asymmetric vortex breakdown features and unfavorable interactions at sideslips are the cause of the limited maneuverability. Means were deduced to control the vortex systems to mitigate or even avoid previous limits. This was done via a combination of theoretical thinking and computational fluid dynamics flow simulations. Aerodynamic means to control the lateral stability were investigated. Wind-tunnel tests secured this process. Small but effective fuselage strakes and leading-edge root extensions between the canard and the wing enhanced the lateral stability considerably. Much higher angle of attack can be reached for increased maneuverability. The design concept was confirmed through intense wind-tunnel campaigns and successfully flown in a test program covering the envelope.

Quasi-steady doublet-lattice correction for aerodynamic gust response prediction in attached and separated transonic flow

A quasi-steady doublet-lattice correction method is used to predict aerodynamic gust responses of two different configurations: a swept wing, the so-called Aerostabil wing, and a transport aircraft configuration, the NASA Common Research Model. The results of the correction method are compared to uncorrected doublet-lattice results, and to results obtained from a nonlinear computational fluid dynamics solver, the DLR TAU-Code. The correction method agrees well with time-marching results obtained by TAU in the limit of dynamically linear gust amplitudes and improves with gust length. In separated transonic flow, an oscillation of the aerodynamic gust response can be computed.

Longitudinal Aerodynamic Performance of the Apollo Entry Capsule near Transonic Speeds

A complex dynamic behavior exists in the flow structures surrounding blunt atmospheric entry capsules as they decelerate from supersonic to subsonic speeds during a lifting reentry. In the current study, unsteady simulations of the Apollo capsule undergoing pitch oscillations were carried out applying the DLR (German Aerospace Center) TAU code to predict the static and dynamic aerodynamic derivatives. Numerical results for static analysis show that the capsule, with its given center of gravity, is predicted to be statically stable both in the super- and subsonic regimes. The dynamic analysis was then performed, indicating that the capsule will turn dynamically unstable when the trim angle of attack is close to zero or if the freestream Mach number decreases. The unsteadiness of wake vortices and the influence of wake pressure on the moment coefficient appear to be the primary factors affecting the dynamic behavior. A time lag between the capsule oscillations and the moments acting on the backcone is shown to be the source of the observed dynamic instabilities.

High-Angle-of-Attack F-16XL Flight Simulations at Sub- and Transonic Speeds

Modern combat aircraft’s highly or moderately swept wings, being designed for supersonic cruise, also have to allow for high maneuverability at trans- and subsonic speeds. This calls for the exploitation of separated flows such as vortices, vortex systems, and their combination with shock systems to the better of the aerodynamic characteristics. In-depth understanding of these flows is vital for the successful design. Here, sub- and transonic free-flight high-angle-of-attack flow simulation results are shown in detail for the F-16XL combat aircraft. Unstructured meshes were adapted manually according to the expected flowfields, covering the complete aircraft without any geometric simplification for the free-flight conditions. The scope of this paper probes for the quality and quantity of high-resolution numerical simulations possible with the current computer capacity at free-flight Reynolds numbers. The DLR Tau code was used to solve the Navier–Stokes equations. The F-16XL (unsteady) Reynolds-averaged Navier–Stokes simulations are performed with the scale-adaptive/detached-eddy simulation turbulence model. Aerodynamic high-angle-of-attack flow features are discussed together with capabilities of the current modeling capabilities and the need for better understanding of vortex system and shock–vortex interference effects.