Benedikt Roidl

Particle-Image Velocimetry and Force Measurements of Leading-Edge Serrations on Owl-Based Wing Models

High-resolution Particle-Image Velocimetry (PIV) and time-resolved force measurements were performed to analyze the impact of the comb-like structure on the leading edge of barn owl wings on the flow field and overall aerodynamic performance. The Reynolds number was varied in the range of 40,000 to 120,000 and the range of angle of attack was 0° to 6° for the PIV and −15° to +20° for the force measurements to cover the full flight envelope of the owl. As a reference, a wind-tunnel model which possessed a geometry based on the shape of a typical barn owl wing without any owl-specific adaptations was built, and measurements were performed in the aforementioned Reynolds number and angle of attack range. This clean wing model shows a separation bubble in the distal part of the wing at higher angles of attack. Two types of comb-like structures, i.e., artificial serrations, were manufactured to model the owl’s leading edge with respect to its length, thickness, and material properties. The artificial structures were able to reduce the size of the separation region and additionally cause a more uniform size of the vortical structures shed by the separation bubble within the Reynolds number range investigated, resulting in stable gliding flight independent of the flight velocity. However, due to increased drag coefficients in conjunction with similar lift coefficients, the overall aerodynamic performance, i.e., lift-to-drag ratio is reduced for the serrated models. Nevertheless, especially at lower Reynolds numbers the stabilizing effect of the uniform vortex size outperforms the lower aerodynamic performance.

Hydrodynamic instability and shear layer effects in turbulent premixed combustion

A turbulent premixed plane jet flame is analyzed by large-eddy simulations. The analysis shows that the flame front wrinkling is strongly influenced by the shear layer effect when the gas expansion effects are small leading to larger flame front amplitudes at the flame base than at high gas expansion ratios. However, the hydrodynamicinstabilityeffect induces a continuously increasing flame front amplitude which yields an enhanced flame pocket generation at the flame tip. Both phenomena influence the magnitude of the turbulent burning area and burning area rate response through the flame front deflections which are determined by the contribution coefficient. This coefficient represents the mutual interaction between the flame and the flow. At low gas expansion ratios, the total heat release rate spectra of the turbulentflame are wider in terms of dominant modes at Strouhal numbers which are linked to the mean flame height oscillations. Thus, at low gas expansion ratios, the vortex-flame interaction is less damped by the flame in the sense that vortices can perturb the flame front stronger. The total heat release rate trend of St−2.2 previously found for a round jet flame is also determined for the current slot jet at realistic gas expansion ratios indicating a general tendency to transfer energy from large to small flame structures. At high gas expansion ratios, an increasing Markstein length leads to an energy transfer between neighboring dominant modes in the low frequency range 1 < St < 10 and the burning area rate response becomes more important for the total heat release rate spectra of the turbulent slot flames which agrees with recent findings for a laminar premixed plane flame.

Particle-image velocimetry investigation of the fluid-structure interaction mechanisms of a natural owl wing

The increasing interest in the development of small flying air vehicles has given rise to a strong need to thoroughly understand low-speed aerodynamics. The barn owl is a well-known example of a biological system that possesses a high level of adaptation to its habitat and as such can inspire future small-scale air vehicle design. The combination of the owl-specific wing geometry and plumage adaptations with the flexibility of the wing structure yields a highly complex flow field, still enabling the owl to perform stable and at the same time silent low-speed gliding flight. To investigate the effects leading to such a characteristic flight, time-resolved stereoscopic particle-image velocimetry (TR-SPIV) measurements are performed on a prepared natural owl wing in a range of angles of attack 0° ≤ α ≤ 6° and Reynolds numbers 40,000 ≤ Re(c) ≤ 120,000 based on the chord length at a position located at 30% of the halfspan from the owls body. The flow field does not show any flow separation on the suction side, whereas flow separation is found on the pressure side for all investigated cases. The flow field on the pressure side is characterized by large-scale vortices which interact with the flexible wing structure. The good agreement of the shedding frequency of the pressure side vortices with the frequency of the trailing-edge deflection indicates that the structural deformation is induced by the flow field on the pressure side. Additionally, the reduction of the time-averaged mean wing curvature at high Reynolds numbers indicates a passive lift-control mechanism that provides constant lift in the entire flight envelope of the owl.

Impact of transversal traveling surface waves in a non-zero pressure gradient turbulent boundary layer flow

The impact of non-zero pressure gradient flow in a turbulent boundary layer flow over a surface undergoing spanwise transversal traveling waves is investigated via large-eddy simulations. While it is known that in zero-pressure gradient flow spanwise surface waves can lead to drag reduction, this question still remains open for non-zero pressure gradient flows. In the present analysis, the effect of a linear pressure gradient is investigated and compared to the zero-pressure gradient flow at a momentum thickness based Reynolds number R e ? = 2000 for a constant surface wave, i.e., the wave length is λ + = 500 , the amplitude A + = 50 , and the wave speed is c + = 6.25 . The results show a drag reduction of about 10% for the zero-pressure gradient flow, 6% for the adverse-pressure gradient, and a drag increase of 4% for the favorable-pressure gradient flow. The analysis of the velocity profiles shows a reduced gradient at the trough region for all actuated setups. At the crest, an increased gradient is obtained. Furthermore, the viscous sublayer is extended. The streamwise turbulent intensity is reduced for all configurations compared to the non-actuated reference case at the crest. At the trough, the shift off the wall is only present for the zero-pressure gradient flow and the adverse pressure gradient flow. The hypothesis based on numerous zero-pressure gradient flow investigations of a reduced wall-normal vorticity component at the crest and trough indicating drag reduction is corroborated. That is, for the adverse-pressure gradient flow the wall-normal component distribution is lowered and for the favorable pressure gradient flow, which possesses a drag increase, the distribution at the trough is similar to that of the reference non-actuated case.

A numerical analysis to evaluate Betz's Law for vertical axis wind turbines

The upper limit for the energy conversion rate of horizontal axis wind turbines (HAWT) is known as the Betz limit. Often this limit is also applied to vertical axis wind turbines (VAWT). However, a literature review reveals that early analytical and recent numerical approaches predicted values for the maximum power output of VAWTs close to or even higher than the Betz limit. Thus, it can be questioned whether the application of Betzs Law to VAWTs is justified. To answer this question, the current approach combines a free vortex model with a 2D inviscid panel code to represent the flow field of a generic VAWT. To ensure the validity of the model, an active blade pitch control system is used to avoid flow separation. An optimal pitch curve avoiding flow separation is determined for one specific turbine configuration by applying an evolutionary algorithm. The analysis yields a net power output that is slightly (≈6%) above the Betz limit. Besides the numerical result of an increased energy conversion rate, especially the identification of two physical power increasing mechanisms shows, that the application of Betzs Law to VAWTs is not justified.

Analysis of spatio-temporal wake modes of space launchers at transonic flow

To detect characteristic wake flow modes responsible for the buffeting phenomenon of space launchers, an optimized dynamic mode decomposition (DMD) combined with a classical statistical analysis is applied to the turbulent wake flow fields of a planar and an axisymmetric generic configuration of an Ariane 5-like launcher at transonic freestream conditions (Ma∞ = 0.8 and ReD = 6 · 10) computed via a zonal RANS-LES method. The investigated wake topologies are characterized by a highly unsteady behavior of the shear layer shedding from the forebody and subsequently reattaching on the nozzle contour. In both cases, the reattachment position features strong oscillations in the streamwise and the spanwise direction, which leads due to wall pressure oscillations to structural loads. The spectral analysis of pressure perturbations on the after-body near the mean reattachment position reveals one dominant characteristic frequency for each configuration, i.e., SrD ≈ 0.23 in the planar and SrD ≈ 0.17 in the axisymmetric case. Moreover, the instantaneous skin-friction coefficient distributions on the nozzle extensions of both configurations indicate the existence of regular longitudinal wedge-shaped patterns with an approximate spanwise/circumferential wave length of 2h in the planar and 120◦ in the axisymmetric case. To clarify the underlying coherent fluid motion responsible for the detected oscillatory behavior, DMD is applied to the time-resolved three-dimensional velocity field. The frequencies of the extracted DMD modes closely coincide with the characteristic peaks in the wall pressure spectra for both configurations. The analysis of the three-dimensional shape of the DMD modes along with the corresponding mean flow modulation in time reveals that the detected periodic behavior is caused by a coherent longitudinal cross-flapping motion of the free shear layer. In the spanwise/circumferential direction, the identified coherent motion of the free shear layer features the same wavelength of ∼ 2h in the planar and ∼ 120◦ in the axisymmetric case as previously indicated by the skin-friction coefficient distributions.

Numerical Investigation of Shock Wave Boundary-Layer Interaction Using a Zonal RANS-LES Ansatz

In this paper a zonal RANS/LES approach is presented in which the regions with attached boundary-layers are computed via RANS and the regions with separated flows by using LES. The transition from RANS to LES takes place in an overlapping region between the RANS and LES zone. Two different turbulent inflow generation methods coupled with a controlled forcing ansatz are applied which enable a fast and smooth transition from two-dimensional RANS- to the three-dimensional unsteady LES solutions. Both approaches require local Reynolds shear stresses of a RANS solution which is located upstream of the LES. The inflow generation methods are validated for a boundary-layer flow and the fully coupled zonal approach is applied to a transonic flow over an airfoil both including a shock boundary-layer interaction.

Zonal RANS-LES Computation for Near-Stall-Airfoil Flow

A fully coupled zonal Reynolds-averaged Navier-Stokes - large-eddy simulation method (zonal RANS-LES) is applied to subsonic flow over the HGR-01 airfoil at high angle of attack. Several zonal-boundary formulations such as forcing layers, a reformulated synthetic turbulent generation method, and a turbulent reconstruction approach are discussed. It is shown that the aerodynamic properties are satisfactorily predicted and the computational costs for a subsonic airfoil near stall compared to a pure LES are decreased by a factor of approximately four. Nevertheless, it has to be stated that the local RANS solution has a non-negligible impact on the susceptible flow phenomena such as the separation when the RANS-LES boundary is located in a non-zero pressure gradient flow regime. This RANS impact is even more pronounced when a pressure driven coupling between an LES-RANS and a RANS-LES boundary exists.

Synthetic Turbulence Generation for a Zonal RANS-LES Method

A synthetic turbulence generation (STG) method for flows at low and high Reynolds and Mach numbers to provide LES inflow boundary conditions of zonal Reynolds-averaged Navier-Stokes (RANS)- large-eddy simulation (LES) method simulations is presented. The present method separates the LES inflow plane into three sections where a local velocity signal is decomposed from the turbulent flow properties of the upstream RANS solution. Depending on the wall-normal position in the boundary layer the local flow Reynolds and Mach number specific time, length, and velocity scales with different vorticity content are imposed on the LES inflow plane. The STG method is assessed by comparing the resulting skin-friction, velocity, and Reynolds-stress distributions of zonal RANS-LES simulations of subsonic and supersonic flat plate flows with available pure LES, DNS, and experimental data. It is shown that for the presented flow cases a satisfactory agreement within a short RANS-to-LES transition of two boundary-layer thicknesses is obtained.

Simulation of Transonic Airfoil Flow Using a Zonal RANS-LES Method

This paper presents a method for a synthetic turbulence generation (STG) to be used in a segregated hybrid Reynolds-averaged Navier-Stokes (RANS)-Large-Eddy Simulation (LES) approach. The present method separates the LES inflow plane into three sections where a local velocity signal is decomposed from the turbulent flow properties of the upstream RANS solution. Depending on the wall-normal position in the boundary layer, the local flow Reynolds and Mach number specific time, length and velocity scales with different vorticity contents are imposed on the LES inflow plane. The STG method is assessed by comparing the resulting skin-friction, velocity and Reynolds-stress distributions of zonal RANS-LES simulations of flat plate boundary layers with available pure LES, DNS, and experimental data. It is shown that for the presented flow cases a satisfying agreement within a short RANS-to-LES transition of two boundary-layer thicknesses is obtained. The method is further used for the simulation of a shock-boundary-layer interaction around an airfoil at transonic flow conditions, where the separated flow region are analyzed by an embedded LES and the remaining flow is determined by a RANS solution.