New Results in Numerical and Experimental Fluid Mechanics XIII

This work describes the influence of the wind tunnel support system on the wake flow and associated forces acting on the CRM aircraft configuration in post-stall conditions at subsonic Mach numbers. Unsteady scale-resolving simulations at $$\alpha =18^{\circ }$$ α = 18 ∘ , $$M_{\infty }=0.25$$ M ∞ = 0.25 and $$Re_{\infty } = 11.6\cdot 10^6$$ R e ∞ = 11.6 · 10 6 were carried out using the CRM with and without a support sting, respectively. A predominantly local effect of the sting was observed at these conditions, causing significant deflection of the local flow direction near the sting. The sting’s shape causes downward deflection near the horizontal tailplane, decreasing that surface’s effective angle of attack and its lift coefficient. This results in a reduced overall lift coefficient $$C_L$$ C L and a significantly reduced nose-down pitching moment $$C_{My}$$ C My . The wing pressure distribution is only weakly affected.

Unsteady CFD simulations of vertical “1-cos” gusts interacting with the DLR-F15 airfoil at transonic flow conditions are analyzed. The scope of the work is to assess the applicability of the simplified Disturbance Velocity Approach (DVA) for gust simulation, which is implemented in the CFD code TAU. While the DVA covers the interaction of the gust with the airfoil, it neglects the effect of the airfoil and the flow around the airfoil on the gust. Simulations where the gust and all interactions are resolved within the flow field are used to evaluate the accuracy of the DVA. Gust wavelength and angle of attack are varied to ensure a reliable assessment of the DVA. Lift, pitching moment, and shock position can be accurately determined by the DVA. At higher angles of attack and a small gust wavelength, the DVA’s drag history shows deviations which should be taken into account. Overall, the DVA provides suitable results with respect to the resolved gust simulations.

Four different methods for the generation of input data for propeller models (Actuator Disk, Actuator Line) in CFD simulations are described and their results are compared with an unsteady fully resolved simulation of the propeller geometry. It is shown that the neglect of the rotation effects with a 2d method overestimates the propeller thrust and the axial velocity in the wake. If the rotation effects are instead considered with a 3d or a new 2.5d approach or via a recovery from the slipstream data, there is better agreement regarding thrust and slipstream velocities. Thereby, the 2.5d method shows a good compromise between computational effort and the quality of the results.

In the field of turbulence research, many experiments are performed using turbulence generated under reproducible conditions. In the past mostly static grids having a fixed blockage were adopted, but in order to generate turbulence with high intensity and high turbulence Reynolds numbers, so called active grids can be used. Therefore, an experimental study has been carried out to investigate the effect of the motion pattern on the turbulence generated by an active grid with individually controllable paddles in the wind tunnel. Measurements were performed using the hot-wire technique. In comparison to former experiments, modifications of the active grid had been made concerning the arrangement of the paddles. The results show that the homogeneity of the turbulence could be improved using the modified grid. Turbulence Reynolds numbers could reach values up to 540. In order to produce even higher integral time scales or higher turbulence Reynolds numbers, synchronous movement of adjacent paddles was tested but did not show the intended improvements. In this paper, the results of the experimental study are presented in more detail.

Numerical simulations of passive scalars in turbulent channel flows up to friction Reynolds number $$Re_\tau =5{,}200$$ R e τ = 5 , 200 and Schmidt number $$Sc=2{,}000$$ S c = 2 , 000 are performed by utilizing the stochastic one-dimensional turbulence (ODT) model as stand-alone tool. The model is calibrated once for the turbulent velocity boundary layer at $$Re_\tau =5{,}200$$ R e τ = 5 , 200 so that the passive scalar is a model prediction. ODT is able to reproduce with reasonable accuracy the scaling regimes of the scalar transfer and locally resolve the boundary layer structure. Albeit the model is unable to capture the emerging dissimilarity of near-wall scalar and momentum transport for high Sc, it can economically and accurately represent fluctuating wall-normal fluxes.

We present an analysis of RANS turbulence models for a new turbulent wake flow case in adverse pressure gradient. The reference data are obtained from an LES. RANS simulations using the Spalart-Allmaras model, the SST model and the SSG/LRR- $$\omega $$ ω model show a tendency to underpredict the onset of flow reversal in the wake. The comparison with the LES data reveals that the SSG/LRR- $$\omega $$ ω model underpredicts the turbulent transport of the Reynolds stresses and overpredicts the ratio of production to dissipation. Therefore two modifications of the SSG/LRR- $$\omega $$ ω model are studied, i.e., the modification of the coefficient of the gradient diffusion hypothesis for the turbulent transport of the Reynolds stresses and the sensitization of the dissipation rate to irrotational strains. Both modifications improve the agreement with the reference data.

Recently, a novel approach to study the inner-outer interactions in turbulent wall-bounded flows has been presented by Mäteling et al. [8], which tackles the combined effects of superposition, bursting events, and amplitude modulation based on the time-dependent streamwise velocity fluctuations. This approach is extended in the current study to all three velocity fluctuation components. Due to the incorporation of the wall-normal velocity fluctuations, the bursting events can be study in more detail and independently of the superposition. By using a multivariate empirical mode decomposition for the scale separation, all velocity components at both wall-normal positions can be decomposed simultaneously to ensure proper scale alignment between the resulting modal representations. The results show that the superposition effect and the amplitude modulation exist in all three velocity components but are most intense for the streamwise fluctuations. A high level of coherence of the sweeps and ejections is also detected by conventional cross-correlation.

Correcting functional terms using data driven techniques is one recent approach to improve the predictive accuracy of turbulence models. We present a multitask learning framework that can be used to generalize functional corrections of turbulence models to different flow conditions and test cases. The approach is developed with a view to generalize the correction of turbulence models using data from different flow conditions. The machine learning model first learns a compressed latent representation of flow variables that are most correlated with the correction term and has enhanced predictive accuracy for new test cases. The approach is shown to improve the performance of Spalart-Allmaras turbulence model in massively separated flow regimes.

Convection velocities of small- and large-scale patterns in a turbulent boundary layer obtained via Direct Numerical Simulation are computed by correlating finite-size interrogation windows between consecutive time steps, which is common practice in Particle-Image Velocimetry. A particular region in the logarithmic layer and the lower defect layer of the mean flow is identified where different patterns belonging to different flow variables are convected downstream by the local fluid material. The corresponding structures are unique and match with the well-known high- and low-speed streaks of the streamwise velocity component. Outside of the region where these structures are observed, the convection of local disturbance patterns exhibits wave character, i.e., they travel faster than the mean flow close to the wall and slower in the outer defect layer and at the edge of the boundary layer. This further confirms our interpretation of the presence of dominant material-bound coherent structures in the intersection zone of the log layer with the defect layer.

This paper presents HyperCODA, the hypersonics extension to the flow solver CODA (CFD for ONERA, DLR and Airbus). Similar to the spacecraft extensions of TAU, HyperCODA extends CODA for applications at high Mach numbers, including non-ideal gas thermodynamics, gas mixtures, and chemistry. The paper discusses simulations demonstrating numerical convergence and stability properties, as well as a supersonic rocket retropropulsion maneuver. Cross-code comparisons are made against the established DLR TAU code.

Experimental determination of NO and $$H_2O$$ H 2 O production during hydrogen combustion applying laser absorption spectroscopy will be performed in the High Enthalpy Shock Tunnel Göttingen of the German Aerospace Center. The combustion process forming these species within the propulsion unit of a small scale wind tunnel model is to be studied. A novel experimental approach to obtain absorption spectra of both species in the infrared region within $$\mu $$ μ -second range is reported. This technique is based on the use of frequency combs generated with quantum cascade lasers. This method enables multi-species measurements of absorption spectra from which concentrations can be inferred. The generated frequency combs are within the $$1760\,\pm \,30\,\mathrm cm^{-1}$$ 1760 ± 30 c m - 1 wavenumber region which features strong NO and $$H_2O$$ H 2 O spectral lines allowing characterization of these species. The technique, the application in short duration ground testing and its calibration procedure will be reported.

The paper will discuss an experimental approach to determine wall surface heat flux in a hydrogen combustion chamber using a temperature sensitive paint. This measurement technique represents a non-intrusive method to determine wall surface heat flux with a high spatial resolution. The application of this method to an internal flow, generated in a hydrogen fueled combustion wind tunnel model, is a new approach in the application of the measurement technique. The wind tunnel model represents the small scaled flight experiment configuration of the MR2 vehicle of the EU co-funded Long-Term Advanced Propulsion Concepts and Technologies project. The paper will discuss the special demands and prerequisites needed to apply TSP in high temperature flows. The experimental setup will be described and the feasibility of the application of TSP to internal flows and the resulting technical demands will be reviewed. The obtained result will be discussed with respect to the validity by comparison to a numerical simulation of the combustor flow.

Aerodynamic characterization of vehicles is mandatory to identify their flight envelope, their controllability behavior and ultimately their shape. During the design process one goal is to efficiently determine the aerodynamic properties for every possible state during the flight. Therefore, the aerodynamic force and moment coefficients need to be known for all combinations that are expected to occur of velocity, orientation, altitude, control surface deflections and dynamic values such as rotation rates. The amount of conditions that need to be evaluated can be reduced if an independence within these variation parameters is proven. For an aerodynamic design process via CFD this paper purposes an additional method to reduce the amount of data points. Through the use of the DLR TAU code with an overlapping meshes feature, envelope datasets can be produced for the momentum free trimmed state in each flight point. With this starting set, several subsets with parameter variations can be computed. This reduces the amount of simulations needed by 65% for the investigated case, the flight experiment ReFEx.

ReFEx is a Reusable Flight Experiment of the DLR. For the layout of the aerodynamic shape of the vehicle, the identification of a flight trajectory and finally the control systems, extensive data sets are necessary. The focus of this paper lays on the comparison of calculated versus measured aerodynamic coefficients. All simulations were carried out with the DRL TAU Code. The measurements were performed in the trisonic wind tunnel (TMK) of the DLR. One aerodynamic flight control element of ReFEx are the canards. The here presented analysis delivers a good agreement comparing the calculated and measured coefficients, looking at the symmetrical canard deflections. The occurred differences of the asymmetrical canard deflection are discussed in detail, including the numerical as well as experimental uncertainties.

The lift reduction potential of fluidic flow control devices is investigated. The devices target unsteady gust load alleviation to achieve a reduction in overall wing weight and fuel consumption. The study focuses on a jet normal to the surface and jet over a Coanda surface installed at the trailing edge. The fluidic actuators are compared to a trailing edge flap for reference using unsteady Reynolds-averaged Navier–Stokes simulations. Unsteady interactions of the airfoil with different gusts are simulated to estimate the required load change and response time of the control devices. To reduce computational costs, a surrogate model based on precomputed aerodynamic polars is developed to estimate gust-induced load changes. The effect of jet temperature for the fluidic actuators and dual blowing for Coanda actuator are studied.

A laminar wing test case for the flutter behavior in transonic flow with free boundary layer transition is presented. The DLR-AE-L1 airfoil and wing are designed to serve as an aeroelastic testbed to develop a better understanding for the effect of boundary layer transition on flutter stability. Coupled CFD-CSM simulations are used to determine the flutter onset in a transitional and a fully turbulent flow. It is found that flutter occurs at lower stagnation pressures for the transitional flow, which is in accordance with earlier research on laminar airfoils.

Hybrid laminar flow control (HLFC) is a promising concept to reduce the viscous drag on aircraft wings by shifting the expected laminar-turbulent transition (LTT) location downstream. However, the structural joint between the leading edge suction panel and the wing box can diminish this desirable shift in LTT location and thus reduce the efficiency of the HLFC system. This paper introduces a toolchain, which allows to numerically quantify the detrimental effects of two-dimensional surface irregularities for transonic HLFC swept wings at free-flight Reynolds numbers. A backward-facing step and two different gaps are studied, which resemble possible structural joint geometries currently considered within the HLFC-WIN project. The stability analysis performed with the Adaptive Harmonic Linearized Navier-Stokes (AHLNS) approach reveals that none of the three irregularities affects the expected LTT location, suggesting that there might be some potential to relax the surface tolerances initially specified in the HLFC-WIN project using classical criteria.

Promising results are obtained by NASA with the CRM-NLF laminar wing at design conditions. The work presented seeks to develop an understanding of the transition mechanisms involved with the CRM-NLF design under wind tunnel conditions. In addition, the capabilities of DLR tools for RANS-based transition prediction are demonstrated. Particular attention is paid to the effect of turbulent wedges. Numerical results are compared to wind tunnel data made available by NASA.

Temperature sensitive paint images from a wind tunnel campaign for the CRM-NLF geometry are used to extract experimental transition lines to be used to determine critical N-factors based on incompressible and compressible linear stability theory. Based on the eN-method, transition locations are subsequently predicted for different flow conditions using RANS computations and the results are compared to the experiment.

The impact of freestream turbulence (FST) on a roughness-perturbed, laminar boundary layer is investigated in a laminar water channel. Both wire-mesh grid and rising air bubbles are used to induce random disturbances to the flow. The transition Reynolds number depends significantly on the level of random fluctuations. This confirms that the lower transition limit, which is relevant to most engineering applications, depends on external noise and is thus difficult to predict without precise knowledge of related flow parameters. Moreover, visualizations and harmonic forcing with a vibrating wire suggest that the near and far wake may respond differently to freestream turbulence. The near wake can effectively be forced in a region at the height of the roughness, whereas the far wake is less sensitive to this type of forcing.

We describe a simple pre-design method to estimate the size of an air outlet flap for a passive hybrid laminar flow control (HLFC) system to determine the feasibility of that flap. The method is based on an ingenious application of old experimental results. The computational effort is minimal.

Suction through porous aircraft wings is a promising concept to reduce its drag, but the modified boundary layer can be more susceptible to external forcing, like surface roughness and inhomogeneities in the suction distribution. This paper analyzes the effects of suction on the linear receptivity of a swept-wing boundary layer to surface roughness and non-uniform suction, focusing on stationary crossflow instabilities. Employing a compressible adjoint approach, receptivity coefficients for a Falkner-Skan-Cooke similarity solution and a transonic swept-wing boundary layer on an A320 vertical tail plane with and without suction are given.

This work presents results of a numerical feasibility study in which HLFC wing designs for a wind tunnel model were obtained for a test in the S1MA wind tunnel. The wind tunnel model will be a testbed for HLFC technology in general and specifically a test of the integrated HLFC technology in the ground-based demonstrator of the European project HLFC-WIN. In the feasibility study a wing should be designed which satisfies the following requirements: 1) Inner wing with high values for attachment transition criteria Reynolds number 2) Outer upper wing surface with cruise design pressure distribution corresponding to the HLFC-WIN ground-based demonstrator 3) Obtained surfaces should allow manufacturability of a metallic microperforated leading edge HLFC panel. The results obtained in this work show that a design is possible which satisfies all three requirements. As a result of the requirements and the short aspect ratio planform of the wind tunnel model wing in comparison to the large aspect ratio of the ground-based demonstrator wing, the wing design led to a wing with large 3D properties. Nevertheless, the designed outer wing satisfies an isobar concept and the surface allows the manufacturability of the leading edge suction panel.

The applicability of Local Stability Theory (LST), Parabolized Stability Equations (PSE) and the Adaptive Harmonic Linearized Navier-Stokes (AHLNS) approach is investigated in the presence of 2-D surface irregularities through comparison with Direct Numerical Simulations (DNS). Remarkably good agreement between DNS and AHLNS is obtained for the amplification curves of Tollmien-Schlichting (TS) waves in all the cases studied. The LST and PSE results exhibit differences which are discussed in relation to the local distortion of the boundary layer induced by the irregularities.

Gust loads on aircraft are critical for structural design. This paper investigates local bump-like modifications in the leading edge region on a generic supercritical airfoil at transonic inflow conditions for the purpose of gust load alleviation. Several parameters are studied to identify a sweet spot for gust load compensation. These include the bump’s height, its streamwise position, its streamwise extension and its oscillation frequency. The evaluations of the URANS simulations reveal that the bumps as investigated in this work are not applicable for compensation of gust induced lift. However, dynamic bumps enable a compensation of approx. 20% of the gust induced pitching moment. A comparison of dynamic bumps with an oscillating trailing edge flap (TEF) and leading edge flap (LEF) reveals the superiority of the flaps for gust load mitigation, especially when a load factor of one is targeted throughout the critical gust event. The qualitative effect of dynamic leading edge bumps on the surrounding flow field is found to be similar to the one caused by oscillating LEF.

Finding fully converged, steady-state solutions of the compressible Reynolds Averaged Navier-Stokes (RANS) equations for aerodynamic configurations on the border of the flight envelope often poses serious challenges to solution algorithms that have proven robust and successful for configurations at cruise conditions. Examples of such cases are agile configurations at high angles of attack. When trying to compute solutions in these scenarios, one often observes that the solution process breaks down after few iterations or that a steady-state RANS solution, although it may exist, cannot be reached with the employed solution algorithm. While, in general, no clear reason for this behavior can be identified, the complexity of these flows seems to be significantly greater compared to flows around transport aircraft in cruise flight. The flow fields are dominated by the interaction of shock waves with a system of vortices emanating from the leading edges on the upper surface of the wing, leading to massive flow separation. These flow features tend to be inherently unsteady and can be assumed to cause problems in computing a converged solution using an algorithm designed to find steady-state solutions of the RANS equations. To avoid these problems, it is not uncommon to calculate such configurations in an unsteady mode, which often comes at a rather high computational cost. This article demonstrates the necessity for implicit smoothers to approximate fully converged solutions of these challenging simulations. A numerical example is given to confirm that convergence is only possible using an exact derivative together with a suited preconditioner.

In this study, experimental investigations are performed to analyze the vortex behavior and breakdown characteristics on a generic multiple swept wing fuselage configuration. Fast Response Aerodynamic Pressure Probe and Particle Image Velocimetry measurements are performed in the low-speed wind tunnel W/T-A facility of the Chair of Aerodynamics and Fluid Mechanics at the Technical University of Munich. The experiments are executed at subsonic speed, symmetric freestream conditions and angles of attack of $$\alpha = 16^\circ , 24^\circ $$ α = 16 ∘ , 24 ∘ and $$32^\circ $$ 32 ∘ . In a power spectral density analysis of velocity fluctuations helical mode instabilities and shear layer instabilities can be detected.

Vortex-dominated flow and their effects on motion-induced air loads in the case of transonic inflow represents a complex and so far not well understood sub-area of aerodynamics. Especially the modeling of such kind of flows with the help of RANS methods represents a major challenge with regard to the choice of initial and boundary conditions. In this article, the results of validation studies are presented using wind tunnel measurement campaigns. The main focus is on the dependence about the angle of attack and the associated characteristics of the vortex-dominated flow.

In the last decades, the interest for light-weight air vehicles and unmanned aerial vehicles (UAVs) equipped with small-scale propellers has risen significantly. Such devices are often equipped with counter-rotating coaxial propellers in order to attain higher thrust levels while maintaining a compact package. In response to their operating scope, these propellers are often subject to a wide range of inflow angles. In this work, a coaxial propeller configuration is investigated under different inflow angles ranging from $$0^\circ $$ 0 ∘ to $$180^\circ $$ 180 ∘ . Wind tunnel experiments are performed to measure the steady loads and to compare the efficiency against two isolated puller propellers. Furthermore, selected flow fields are captured using smoke visualization. Emphasis is on identifying the interactions of the propeller wakes and the vortex structures. In parallel to the experimental study, a series of URANS simulations are performed in order to study the loads and the flow field in a time-resolved manner.

The present work discusses a methodology to reconstruct three-dimensional positions of rotor blade tip vortices based on stereoscopic background-oriented schlieren data. The reconstruction procedure is first applied on a rotor test stand where a comparison with previous PIV measurements is performed to validate the reconstruction scheme. Subsequently, it is adapted for the reconstruction of the vortex system surrounding a model helicopter in free flight during an unsteady takeoff maneuver.

Particle Image Velocimetry (PIV) is an important technique to investigate complex flow fields but limited to the tracking fidelity of the flow tracers and the uncertainties in the evaluation process. To examine these issues this work implements a physic-based motion simulation for small spherical particles, serving as flow tracers in PIV, with existing tools for image generation. Particular focus set on the behavior in vortical structures and the occurrence of “particle voids” due to centrifugal forces. The results show that a realistic particle distribution ideally can be achieved when applying a statistical distribution of the particle size. Furthermore, they suggest a significant error in the velocity prediction if the region of interest has no sufficient particle seeding.

Small-scale coaxial-rotor assemblies are commonly utilized in Unmanned Aerial Vehicles to provide a compact design solution for achieving increased thrust potential and flight stability. The presence of two rotors operating in close proximity to each other results in aerodynamic interactions that have a profound effect on the individual rotor performance. A dedicated coaxial rotor test bench was developed for the purpose of testing various setups in a multi-rotor arm configuration. This study investigates the influence of different propeller blade geometries, number of rotor blades, and rotor separation distances on the aerodynamic, electric, and system performance of coaxial rotor configurations in static thrust conditions. Furthermore, the thrust and power measurements for one configuration is presented in axial and reverse flow conditions.

The modifications to the flow caused by fairings placed outside a cycloidal rotor are studied. Six configurations using these passive aerodynamic surfaces are covered. The rotors perform fully developed hover flight with a 2D URANS CFD model. Two principal flow regimes emerge from the six configurations. One regime reproduces the blade vortex shedding typically found in cycloidal rotors which occurs as the blade travels upwards. The other regime does not exhibit it and the blades instead shed two smaller vortices as they reach the bottom of the rotor. This modified flow regime rotates the thrust vector and the wake by roughly 25 $$^\circ $$ ∘ . It also significantly reduces both power consumption and drag. A rotor with opened fairings on both sides yields a 6% increase of the thrust to power ratio at equal thrust.

The fan boundary condition in Ansys Fluent was applied to a helicopter rotor in vertical flight using a Robinson R22 geometry. This simplified boundary condition, which is based on a pressure jump over an actuator disk, offers considerable advantages in speed and stability compared to methods using a blade element theory. It was shown that the identification of different rotor working states and preparation of the induced velocity curve is possible by applying the analyzed rotor model. Particular attention was given to the prediction of the vortex ring state - a phenomenon that cannot be described using the momentum theory and which poses a significant threat during helicopter flight. The outcomes were comparable to those of experimental visualizations and simulations performed using the more computationally expensive Virtual Blade Model, thereby proving the viability of the fan boundary condition to model the main rotors.

Due to various design variables, like the blade shape and operating conditions, the optimization of propellers is a complex task. The aerodynamic and aeroacoustic design optimization is conducted by using an efficient machine learning optimization approach, the Bayesian optimization. On the one hand, this allows for a fast and efficient optimization process. On the other hand, the knowledge about the optimization problem is extend and the influence of the individual design variables can be identified. Exemplarily, the radial airfoil sections of the propeller blade are optimized regarding the glide ratio and the aeroacoustic footprint.

In order to estimate boundary layer properties and flow separation within the mid-fidelity aerodynamics code UPM, fast, simple, and robust methods are required. Following a review of suitable methods, two approaches were selected and implemented: stripwise analysis using integral boundary layer methods for lifting surfaces and a simplified analysis based on flat-plate analogy for arbitrary non-lifting bodies. The implemented methods were validated for airfoil, wing, fuselage, and rotor simulations. The results show that the implemented methods accomplish the original purpose of estimating friction forces for bodies in attached flows and detecting regions probably dominated by flow separations. However, they neither model detached flows nor predict their impact on body forces and moments.

Besides aerodynamic performance, most unmanned aerial vehicles (UAVs) are limited by noise emission in their range of application. Especially in urban areas, the acceptance of UAV noise emission by the affected population is required. Therefore, one objective of DLR’s project URBAN RESCUE is to investigate the noise emission and noise generation mechanisms of small rotors. In order to study noise emission, a test setup was established in DLR’s 1m wind tunnel. For different flight modes, acoustic measurements were conducted using a microphone array. For all test cases the noise emission of the rotor’s blade passing frequency (BPF) is higher in front of the rotor (upstream) than behind it. Due to higher relative velocities on the advancing side than on the retreating side, emission is asymmetric with respect to the flow direction. Additionally, a time-resolved background-oriented schlieren (BOS) setup was used to visualize tip vortices and study noise generation mechanisms.

The flight dynamics and flight envelope of coaxial ultralight helicopter CoAX 2D were investigated. For this purpose, series of flight tests was carried out. This paper presents the computation of the fuselage aerodynamics of the CoAX 2D with the DLR flow solver TAU. For cruise flight, fuselage polars with different angles of attack and yaw angles without rotor downwash were calculated. Furthermore, different airspeeds, climb and sink rates were simulated in combination with rotor downwash. To simulate the rotor downwash from the coaxial rotors, the Actuator Disc approach is used. The lift of the horizontal stabilizer is compared between flight test and simulations for validation purposes. The lift was determined in the flight test by deflection of the tail boom. Finally, concepts for the optimization of the fuselage were developed. The landing gear was selected as a possible optimization potential. The changes from an aerodynamic point of view related to the shape of the skid supports. The aerodynamics of the new landing gear were compared with the old configuration and a reduction of drag could be shown.

The flight of a helicopter within the wake of a preceding fixed-wing aircraft, for example during air-to-air refueling, is accompanied by vortex-rotor interactions. The full mutuality of vortex and rotor wake interactions requires a free-wake solution that is applied at different advance ratios in this paper and compared to existing simplified models. The results indicate that the simplified models are valid for higher advance ratios whereas very small advance ratios require the use of a free-wake approach.

The requirements for a computational fluid dynamics solver to simulate the vortex-dominated helicopter rotor flows correctly are high. In recent years, many higher order methods and strategies have evolved to meet these requirements. This paper investigates a 3rd and 4th order accurate upwind scheme and compares this with a hybrid approach, where a 2nd order solution is generated near the rotor blades and a 4th order solution is computed in the wake. The investigated test cases are an inviscid vortex, an inviscid NACA0012 computation at zero-lift, two model rotors in hover and one model rotor in descent flight. From the exploration of different combinations of numerical schemes for the inviscid flux computation, it is seen that higher order methods are necessary to capture the trends of helicopter flows correctly. A modification to the upwind scheme is proposed to further decrease numerical dissipation and increase tip-vortex preservation. The outcome is that the 4th order accurate upwind scheme with the added modification prevails over the existing strategy.

The scattering of noise generated by helicopter rotors has been recognized as having a significant influence on both the noise spectra and the noise directivity. This paper presents the research by DLR within the collaborative effort in the GARTEUR Action Group HC/AG-24. Shielding experiments were conducted in DLR’s Acoustic Wind tunnel in Braunschweig (AWB) using point and rotor sources. For the numerical simulation of the various test cases, DLR’s boundary element method was applied. The acoustic scattering predictions are compared to both test data and analytic solution.

With the aim to detect bi-stability in the highly separated flow around a sphere mounted on a cross-stream rod, wind tunnel measurements and numerical simulations with a Lattice Boltzmann method (LBM) are performed and compared. In both, LBM simulation and experiment, a bi-stable flow behaviour is detected at a Reynolds number of $$\mathrm {Re}=10^5$$ Re = 10 5 based on the sphere diameter, the inflow velocity and the kinematic viscosity, whereas such behaviour is absent at Reynolds numbers as high as $$\mathrm {Re}=3\times 10^5$$ Re = 3 × 10 5 . The bi-stable behaviour detected in the time series of the aerodynamic side force acting on the sphere for $$\mathrm {Re}=10^5$$ Re = 10 5 is reflected by a bi-modal probability density function. For $$\mathrm {Re}=3\times 10^5$$ Re = 3 × 10 5 , on the contrary, the probability density function is nearly Gaussian. Further, in agreement with the experiment, two counter-rotating vortex tubes originating from the intersection of the leeward and cross-stream rod-ward side of the sphere are generated.

Blood flow in channels of varying diameters $$< 500\,{\upmu } \mathrm{m}$$ < 500 μ m exhibits strong non-linear effects. Multiphase finite volume approaches are feasible, but still computationally costly. Here, the feasibility of applying convolutional neural networks for blood flow prediction in artificial lungs is investigated. Training targets are precomputed using an Eulerian two-phase approach. To match with experimental data, the interphase drag and lift, as well as intraphase shear-thinning are adapted. A recursively branching regression network and convolution/deconvolution networks with plain skip connections and densely connected skips are investigated. A priori knowledge is incorporated in the loss functional to prevent the network from learning non-physical solutions. Inference from neural networks is approximately six orders of magnitude faster than the classical finite volume approach. Even if resulting in comparably coarse flow fields, the neural network predictions can be used as close to convergence initial solutions greatly accelerating classical flow computations.

Temperature and humidity measurements are conducted in mixed convective humid-air duct flow with condensation. The latent and total heat transfer during the experiment are determined through the thermal balance for inlet temperatures from 27.5 $$^{\circ }$$ ∘ C to 35.5 $$^{\circ }$$ ∘ C, relative humidities from 30% to 55% and at four Reynolds numbers (2000–8000). The experimental results are compared with a heat transfer model from the literature. Adjusted in terms of the geometry and surface properties, the model shows partial agreement for the cases with forced convection but has to be further adjusted regarding the influence of thermal convection.

In this paper, we present a new mobile respiration simulation system (RSS), which can be connected to existing thermal manikins. With the objective to simulate the human respiration process as realistic as possible, the system was validated on the basis of literature data and results obtained from human subject tests. The RSS reproduces realistic respiration cycles characterized by a sine wave of a typical normal breathing flow rate. The provided flow rates as well as the breathing frequency – representing the time of inhalation and exhalation – were verified by literature values. Since the new system additionally allows to enrich the exhaled air with carbon dioxide ( $$\mathrm{CO}_2$$ CO 2 ), experimental studies addressing the indoor air quality are also feasible. Here, the amount of $$\mathrm{CO}_2$$ CO 2 emitted by the RSS corresponds to the average amount of $$\mathrm{CO}_2$$ CO 2 exhaled by test persons. In addition, the flow characteristics occurring in a human nose are simulated using a self-developed facial mask, in combination with the new system. The result is a breathing thermal manikin based on a mobile respiration simulation system, which can easily be connected to heated passenger models. Accordingly, the system can be installed at any seat within a passenger compartment. This offers the advantage of individually defining the location of the manikin, which can effortlessly be adapted during a measurement campaign. Therefore, the system especially suitable for studies addressing the performance of ventilation systems in passenger compartments and indoor environments.

A detailed description of an experimental set-up designed for upcoming investigations of latent and sensible heat transfer in a cuboid sample with air in- and outlets is given. The container is heated at the rear and cooled at the transparent front wall. Temperature measurements reveal that both sides exhibit a mean temperature deviation below 2% relative to the temperature difference between the mean plate temperature and the ambient temperature. This is a suitable temperature distribution for such measurements. Tomographic particle image velocimetry covering the entire volume of the mixed convection cell exhibits a large-scale circulation due to forced convection with a buoyancy-induced flow close to the temperature controlled surfaces. Forced convection origins from a flow between the inlet and the outlet channel with a mean deviation of 1% from the mean velocity and a maximum absolute deviation of 0.04 m/s. Measurements were performed for Reynolds numbers ranging from $$300< \mathrm {Re} < 2000$$ 300 < Re < 2000 and Grashof numbers $$\mathrm {Gr} < 1.2 \times 10^8$$ Gr < 1.2 × 10 8 . A representative flow field obtained at $$\mathrm {Re}=620$$ Re = 620 and $$\mathrm {Gr}=1.1\times 10^8$$ Gr = 1.1 × 10 8 is presented as an example.

Today you can find various standards to assess thermal comfort. However, it lacks on standards for the assessment of thermal comfort in car cabins, particularly in case of high momentum air-flow or draft. Thermal comfort is often perceived differently in cars compared to e.g. in buildings or coaches. With the objective to overcome this problem, an experimental study of the impact of the draft on thermal comfort is carried out. In this paper, we present results of an investigation focusing on the influence of turbulence intensity on thermal comfort. Draft is simulated in a generic car cabin in the following parameter ranges: temperature $$T\,=\,17~^\circ $$ T = 17 ∘ C– $$29~^\circ $$ 29 ∘ C, velocity $$U\,=\,0.25$$ U = 0.25 m/s–2.5 m/s and turbulence intensity $$I_{LT}\,=\,16\%$$ I LT = 16 % and $$I_{HT}\,=\,32\%$$ I HT = 32 % . The thermal comfort is determined by means of a thermal manikin and infrared thermography to evaluate the thresholds of thermal comfort based on the German Industrial Norms [1, 2].

Two computational approaches are taken to characterize the drag of a car driven behind a heavy vehicle under real conditions. The on-road approach uses velocity measurements, obtained from an array of static five-hole probes, to construct on-flow boundary conditions replicating the atmospheric dynamics encountered during an on-road measurement. The wind tunnel approach uses an oscillating flap system to control flow time and length scales upstream of a wind tunnel model. The amplitude and frequency of the flap motion is calibrated to reproduce length and time scales at on-road conditions under respective Reynolds and Strouhal number scaling. These approaches are evaluated against experimental measurements using Computational Fluid Dynamics (CFD) and demonstrates that they reproduce the aerodynamic drag and a significant part of the measured onflow condition for the car.

The transient incoming flow a compact car experiences whilst driving 10 m – 100 m behind a heavy vehicle on a runway has been characterised. The incoming flow was measured using a 2D array of 11 five-hole probes mounted 1 m in front of an operational, full-scale compact car. Additionally, 188 surface pressure taps were used to measure the effect of the incoming flow conditions on the compact car. The experiments were performed under ideal conditions on a 2.9 km long runway in Faßberg near the DLR in Trauen, Germany.The experiments showed that the heavy vehicle and the distance between the two vehicles have a significant impact on the flow characteristics: with decreasing distances between the vehicles the fluctuations increase from ±0.5 m/s to ±6 m/s and exhibit coherent dominant frequencies at 2.5 Hz in the spanwise direction that are attributed to the vortex shedding in the wake of the heavy vehicle. The turbulence intensities increase from 1% to 25% while the length scales decrease from 25 m to 0.25 m. The change in the incoming flow has a demonstrable influence on the surface pressure of the compact car. The pressure coefficient fluctuates by up to ±0.6 and cp, rms increases from 5% to 30%. The effect of the heavy vehicle on the pressure distribution indicates that the global forces acting on the compact car are influenced.

It is assumed, that the presence of nonlinear effects in transonic flow influences the aerodynamic load distribution and hence must be considered in the aeroelastic loads analysis of maneuvers. In this study a steady pull-up and two quasi-static rolling maneuvers are examined using the Reynolds-Averaged-Navier-Stokes (RANS) equations and compared to the vortex lattice method (VLM). In contrast to the RANS equations, the VLM is widely used for fast analysis of many load cases but does not include transonic effects. The pitch and roll rates are accounted for by the flow solver. A fully elastic transport aircraft configuration is analyzed. Elastic deformations are calculated by the modal approach and are loosely coupled with the aerodynamic model.The resulting loads are compared to maneuver analysis based on the VLM. For all maneuvers, the aerodynamic loads based on the RANS equations yield higher loads at the wing’s root by $$1\%$$ 1 % to $$2\%$$ 2 % and require larger negative balancing forces on the tail plane.

Unmanned or micro aerial vehicles are designed to perform well in a variety of flight conditions. Therefore, the morphing wing technology aims to constantly adapt the aerodynamics to different flight stages. This work presents experimental investigations of an elasto-flexible membrane wing at a Reynolds number of 264000. The investigated concept enables wing folding over a wide range and it allows the wing to adapt to changing aerodynamic loads. The present study focuses on the deformation of the elasto-flexible membrane and its advantages for the aerodynamic performance of the examined model. The measurements show an increasing camber and thickness of the airfoil for higher angles of attack, which results in a more smooth and delayed stall, in particular for the highly swept configuration.

Numerical aeroelastic investigations are conducted for the Model53 delta wing with a deployed slat and two trailing-edge control surfaces at transonic speed and high dynamic pressure. The numerical method is based on coupled high-fidelity CFD-CSM simulations, which is implemented in the multi-disciplinary simulation environment SimServer. The DLR Tau Code is utilized to solve the Reynolds-Averaged Navier-Stokes (RANS) equations and a modal solver is employed to calculate the structural displacements. Control surface regions are modeled with a Chimera approach for hybrid grids. The main focus of the analysis is set on the appropriate handling of control surface deflections for coupled CFD-CSM simulations. Two cases are investigated: In the first case, the control surfaces are deflected for both, the aerodynamic and the structural grid. In the second case, only the control surfaces of the aerodynamic grid are deflected and the aerodynamic forces are transferred to the undeflected structural grid. Both cases are compared in terms of the resulting structural deformation and flow field. The differences between both approaches are minor.

This investigation aims to provide a validation framework for the dynamic simulation tool ASTOR (AircraftEngine Simulation for Transient Operation Research) for gas turbines. ASTOR is based on a system of ordinary equations which is able to simulate transient performance and dynamic response. The numerical setup of ASTOR is compared to state of the art transient performance simulation techniques for the validation. It is shown that ASTOR is able to calculate the performance of the single-spool turbojet engine under investigation to a level of detail that is in no way inferior than other approaches. Furthermore, an experimental validation is carried out using the in-house JetCat P200SX engine. Measurements of steady state operating points and transient maneuvers show a high degree of consistency with performance simulations conducted in ASTOR.

In modern Launcher Rocket Engines (LRE), the study of the behavior of the injector in combustion chamber has been complicated by the fact that lower fuel temperature and higher pressure are used to achieve larger enthalpy values. In addition to this, methane has become more and more common in aerospace application, therefore transcritical states are often reached by the mixing of fuel and oxidizer. The goal of this work is to correctly characterize the transcritical mixing of coaxial $$N_2/H_2$$ N 2 / H 2 injector. Real-gas thermodynamic is used for this purpose in combination with appropriate inlet turbulent boundary conditions. In particular, an effort has been made to use realistic inlet turbulent conditions and a high order scheme has been employed for the calculation of the convective fluxes.

Aiming to enable the use of mid-fidelity panel methods over low-fidelity potential methods in the context of mass simulations for aircraft design, a two-dimensional steady-state prototype was implemented to demonstrate the feasibility of panel methods for aerodynamic simulations. In contrast to most state-of-the-art panel codes, this method is based on high-order singularity distributions and a Galerkin formulation. This paper demonstrates several advantages of the method over the more traditional low-order collocation-based methods.

In the present investigation, the unsteady flow field around a standstill offshore wind turbine is analysed in the hovering and approach regions of maintenance helicopters in order to identify critical flight regions. Three different inflow conditions are analysed, at the rated and cut-off wind speeds. One fictive hovering region and two fictive flight regions for maintenance helicopters are defined and the flow field is separately analysed in these regions. The turbulence kinetic energy and the vertical inclination angle are evaluated. These give indirectly information about the expected blade loads and possible stall occurrence, and a turbulence criterion based on the standard deviation of the vertical velocity component is applied in order to identify possible critical flight regions. The analyses based on the turbulence kinetic energy and on the turbulence criterion suggest the presence of critical areas in the wake of the nacelle for the high wind speed only. Moreover, the analysis of the vertical inclination angle suggests high fluctuations of the velocity direction in the nacelle wake even for the rated wind speed, with possible stall occurrence in the rotor inner part.

The discontinuous Galerkin (DG) method is a high-order method in space reducing the amount of cells needed for calculations compared to traditional computational fluid dynamics (CFD) solvers. The present study concerns the implementation of a DG Chimera method and its application to simulations with relative body movements. The chosen test cases of a rotating 2D cylinder and a slowly plunging NACA0012 airfoil deal with body movement and unsteady laminar flow at subsonic Mach numbers. The results obtained agree well with the reference. Furthermore, the flexible application of the DG Chimera method to unstructured grids is demonstrated.

This study investigates the applicability and efficiency potential of analytical wall functions in conjunction with hybrid RANS-LES methods in two unstructured flow solvers, i.e. the compressible DLR-TAU code and the incompressible DLR-THETA code. Wall-modelled LES simulations of a periodic channel flow using IDDES confirm the validity of the wall-function approach over a broad range of grids with $$12.5\le y^+(1)\le 210$$ 12.5 ≤ y + ( 1 ) ≤ 210 . For the more complex separated flows over a backward facing step and a wall-mounted hump, DDES simulations show that $$y^+(1)$$ y + ( 1 ) has to be limited to 25 regardless of the flow solver, in order to achieve agreement with results using full near-wall resolution. Despite that restriction, the computing time can be reduced by up to 2/3 compared to the fully-resolved case.

RANS simulations with the Spalart-Allmaras turbulence model are improved for cases with flow separation using the Field Inversion and Machine Learning approach. A compensatory discrepancy term is introduced into the turbulence model and optimized using high-fidelity reference data from experiments. Influences on the optimization results with respect to regularization, grid resolution and areas in which the optimization is active are investigated. Finally, a neural network is trained and used to augment simulations on a test case.

We introduce Spliss, a novel block sparse linear solver library, which is designed to meet current and upcoming challenges in the CFD development. Spliss supports a wide range of linear operators typically used in CFD applications. This includes sparse block matrices of variable block sizes and different scalar types as well as matrix-free operators. The provided solving methods comprise standard methods such as Jacobi or Gauss-Seidel smoothers, Krylov subspace methods, and domain-specific methods such as direct block-tridiagonal solvers. Spliss uses contemporary and emerging HPC technologies such as one-sided communication, hybrid and heterogeneous parallelization. The C++ template interface allows for extensions of the library and specialization of its internals for optimization. We explain the integration of Spliss in the two CFD solvers CODA and TRACE. Finally, we show how CODA benefits from using Spliss by offloading the linear solver to GPUs.

In the context of finite-volume methods it is generally accepted that at least second order discretization schemes should be considered for the mean flow equations due to their superior accuracy compared to first order schemes. For turbulence model equations, the situation is not so clear when the Boussinesq approximation is applied. Since the coupling of the turbulence model into the mean flow equations is rather weak (via the eddy viscosity and a rather small contribution to the turbulent kinetic energy to the total energy conservation), one can hope for a minor effect on solution accuracy. The goal of this article is to identify the relevant details required in a careful implementation of a first order treatment of the convective part of the turbulence model, such that the obtained results remain comparable to a fully second order discretization. Compelling arguments are given that a first order treatment of turbulence does not necessarily lead to a relevant loss of accuracy in the investigated cases, while yielding a significant gain in robustness.

A cavity configuration features a highly unsteady separated flow, which is characterised by an intense aero-acoustic coupling mechanism. Strong pressure oscillations produced in and around the cavity have the potential to cause structural fatigue and induce resonance phenomena inside the cavity. In this study, a novel open cavity configuration with doors attached on the sides and a length to depth ratio of 5.7 has been studied numerically using the TAU code developed by German Aerospace Center for high subsonic and supersonic flows. The turbulence is resolved partially using a hybrid RANS-LES turbulence model. The study comprises the Mach numbers (Ma) 0.8 and 1.2 with Reynolds number (Re) $$12 \times 10^{6}$$ 12 × 10 6 . The Rossiter modes occurring in the cavity due to the feedback mechanism have been numerically computed and validated for both flow conditions using reference data from experimental measurements.

A hybrid Reynolds-averaged Navier-Stokes (RANS)/large-eddy simulation (LES) approach for flows with porous surfaces is presented. For the two-dimensional geometry considered, a computationally cheap 2D-RANS simulation in the entire domain is performed first, followed by a LES in regions, where more accurate unsteady flow data are required. For the RANS part a two-equation low Reynolds number k- $$\varepsilon $$ ε turbulence model is modified to include the porous treatment. The boundary conditions for the LES, which is fully embedded inside the RANS domain, are derived from the RANS solution. To enforce a physically realistic transition from an averaged RANS solution towards a resolved turbulent flow field, at the inflow of the LES coherent structures are generated by means of the reformulated synthetic turbulence generation (RSTG) method. The methodology is applied to a NACA 0012 airfoil with a solid and porous trailing edge in a subsonic flow. The simulation is set up with the broadband turbulent boundary-layer trailing-edge (TBL-TE) noise prediction as a future objective in mind, i.e., the noise sources in the trailing-edge region are captured by the LES.

High-fidelity flight maneuver simulations are crucial for the development of realistic digital aircraft models. However, such simulations are still hampered by difficulties in modeling the relative body motion between control and lifting surfaces when using realistic configurations. The presence of spanwise gaps between lifting and control surfaces impedes the application of concepts such as mesh deformation, and hampers the usage of mesh deformation combined with the overset method since the mesh generation process is particularly cumbersome. To reduce the user effort to create overset meshes, we have developed a methodology to automatically create overlapping regions for matching block interfaces. Hence, the usage of the overset method combined with mesh deformation for modeling moving control surfaces is facilitated, and a significant advance towards the computation of high-fidelity flight maneuvers is achieved.

A shape optimization approach backed by adjoint RANS is suggested, which allows for helicopter cell-shape design in powered flight conditions and is realizable in context of industrial design. To this end the DLR TAU code was extended to support an actuator-disk rotor modelling for the discrete-adjoint gradient computation. The steady-state approach accounts for rotor effects in the shape optimization and reliefs from the burden associated with time-accurate adjoint computations for transient rotor analyses. The adjoint CFD method was included in a tool-chain for gradient-based helicopter cell-shape optimization in conjunction with a free-form parametrization of the cell shape and a radial-basis function approach that propagates surface deformations into the CFD volume mesh. The suggested approach was successfully demonstrated in a drag minimization study for the modified helicopter fuselage of the ROtor-Body-INteraction (ROBIN) configuration combined with a realistic load distribution for the 7AD main rotor. The shape optimization was conducted for forward flight conditions at full-scale using unstructured grids. The study confirmed the practical value of the suggested approach: accounting for the rotor downwash in the cell-shape design optimization turned out to be a critical factor for success; the cell shape obtained from an optimization in which rotor effects were neglected deteriorated the aerodynamic performance in powered flight conditions.

The application of reduced-order modeling (ROM) techniques in the context of aerodynamic nonlinear system identification of realistic aircraft configurations gained increasing attention in recent years. Therefore, in the present study the application of a recurrent neuro-fuzzy model (NFM) that is serial connected with a multilayer perceptron (MLP) neural network is introduced concerning the computation of transonic buffet aerodynamics. In particular, the intention of the ROM is the prediction of coefficient time-series trends in contrast to a precise resolution of detailed flow effects. Further, a reduction of computational time compared to a full-order reference Computational Fluid Dynamics (CFD) solution is pursued. The training of the ROM is accomplished based on a data set computed by means of unsteady Reynolds-averaged Navier-Stokes (URANS) simulations. The performance of the trained ROM is demonstrated by predicting the buffet flow characteristics of the NASA Common Research Model (CRM) investigated at transonic flow conditions. Therefore, the wing of the configuration is excited by an external pitching motion beyond buffet onset. By comparing the ROM result with a reference URANS solution, a precise prediction capability of the aerodynamic characteristics as well as a reduction in computational time is demonstrated.

The aerodynamic behavior inside a line-cavity system is investigated within this work. Acoustic effects, like attenuation and resonance, are mainly dependent on the geometric line-cavity system properties: radius r, line length L and cavity volume V. In order to determine the transfer function from the system entry to the location of the pressure sensor at the cavity end, newly developed fiber-optic differential pressure sensors are used to acquire signals of high bandwidth. In contrast to approaches in the frequency domain, where e.g. a speaker emits signals of dedicated frequencies, in this work, the transfer function is calculated in the time domain. A step pressure change in a shock tube is produced and leads to the excitation of frequencies in a large bandwidth simultaneously. In addition to the fiber-optic pressure sensor at the end of the line-cavity system, a further fiber-optic sensor is flush mounted to the shock tube test section as a reference. By applying system-identification routines, the transfer function can be deduced. Experimental investigations of two line-cavity systems of various lengths show very good results. The signals of the reference pressure signals can be reproduced very accurately.

The wind tunnel of the faculty of mechanical engineering at the Unviversität der Bundeswehr München is used for the investigation of two-dimensional airfoil characteristics. The measurement results in the open test section showed a relatively low repeatability and additionally larger deviations from reference data after the application of test section interference corrections. Therefore, the testing procedure is improved and the interference corrections are changed to a method, which represents the actual test situation much better. As a result, the measurements are very repeatable and the corrected airfoil polars match very well with reference data from larger wind tunnels and XFOIL calculations.

Horizontal Axis Wind Turbines are operated with different safety mechanisms which protect the wind turbines from overloading during strong wind situations. However, wind turbines may collapse due to sudden gust events or other overload situations. An additional safety measure may be an aerodynamic load limit of the airfoils themselves by adding passive lift limiting devices to the wing. Ideally, these devices would not add extra drag and the aerodynamic performance will maintain in the normal operation range of the wings. The investigated small radii or ‘noses’ added to the leading edge of a laminar airfoil shows very promising results: the lift limit can be adjusted by the shape, position and spanwise distribution of the added ‘nose’. Additional drag can be kept to a minimum within a certain operation range of the airfoil, depending on the geometrical characteristics of the added nose.

The aim of this work is to impose a disturbance on the velocity field already in the near field of the wake vortex system of a generic high-lift transport aircraft configuration by oscillating flaps. This should result in an excitation of long wave instabilities in the downstream development of the wake vortex system. For aerodynamic force measurements and particle image velocimetry measurements, experiments are carried out in the Göttingen-type low-speed wind tunnel A of the Chair of Aerodynamics and Fluid Mechanics of the Technical University of Munich. Complimentary to the experimental data, unsteady Reynolds-averaged Navier–Stokes flow simulations are conducted. A baseline configuration with statically deflected flaps is compared to a configuration with actuated flaps. It can be shown experimentally as well as numerically that a significant disturbance can be imposed already in the near field of the wake vortex system by oscillating flaps.

A fast and physical-principles-based method to compute the noise radiation for installed propellers’ applications is introduced. The simulation of the sound propagation includes newly defined perturbation equations, with propeller sound sources modelled as rotating singularities. The system of equations proposed is derived from the Linearised Euler Equations, after the application of a split-approach on the velocity perturbations. Two separate coupled systems of equations are obtained, the Acoustic Perturbation Equations, governing the acoustic velocity and pressure perturbations, and the Vortical Convection Equations, describing the convection of non-acoustic velocity perturbations. The moving sources are regularised with a Gaussian kernel over the computational mesh, allowing their description within an unstructured quadrature-free Discontinuous Galerkin experimental solver for Computational Aeroacoustics applications. The prediction capabilities of the method proposed have been successfully validated with a simple test case.

A round jet with $$Re=3600$$ R e = 3600 and $$M=0.9$$ M = 0.9 as well as two planar nozzle-wing flap configurations were simulated by using LES. An evaluation was proposed to predict the effect of the instantaneous Lighthill sources, S, their linear components, $$S^{\mathrm {l}}$$ S l , and non-linear components, $$S^{\mathrm {n}}$$ S n . The simulations indicate that conclusions can be made regarding the directional characteristic of the emitted sound field. These findings were used in investigations of a nozzle-wing-flap configuration. The wing-nozzle configuration showed increased values of the mean Lighthill source terms on the bottom of the wing and flap, contrary to the case of an isolated wing.

A far-field solver based on the acoustic analogy of Ffowcs Williams and Hawkings for acoustic problems with convection and vortical waves was implemented in the open source package Overture. The numerical calculations of CAA test cases showed good agreement with analytical solutions or results from measurements and simulation data from the literature. This paper presents the implemented solver and its verification. In futher work, the developed framework will be used to investigate jet-flap interaction and the emitted sound field, numerically.