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Probst A.,TU Braunschweig | Probst A.,Institute of Fluid Mechanics | Radespiel R.,TU Braunschweig | Radespiel R.,Institute of Fluid Mechanics | And 2 more authors.
AIAA Journal | Year: 2012

An extended eN-based modeling approach for Tollmien-Schlichting-type transition in aerodynamic flow simulations with the low-Re "h-Reynolds- stress model is presented. Instead of simply activating the turbulence production terms at the transition location (point-transition approach), the method incorporates the otherwise neglected Reynolds-stress contributions by the fluctuations of the Tollmien-Schlichting waves and provides them as local input for the turbulence model at the transition point. The shapes of the Reynolds-stress profiles are derived from linear stability analysis within the eN method, whereas their absolute magnitudes are calibrated with the aid of direct numerical simulation data of a transitional boundary layer with adverse-pressure gradient. The dissipationrate input is adjusted to theoretically match the amplification rate of the fluctuations but requires a correction to account for the low-Re damping in the "h-Reynolds-stress model. The paper describes the general modeling ideas and the implementation in the flow solver. Aspects of the numerical discretization and a verification for threedimensional flows are addressed as well. Besides a basic validation for the adverse-pressure-gradient boundary layer, simulations of the SD7003 airfoil flow comprising a laminar separation bubble are presented, which yield very good agreement with measurements. Results of a transitional flat-plate flow are, however, impaired by the lack of intermittency modeling. Finally, the method is applied to a flowthrough nacelle near stall conditions in order to prove its ability to compute consistent transitional behavior in complex three-dimensional flows. Copyright © 2011 by the authors.

Hempert F.,Robert Bosch GmbH | Hoffmann M.,Institute of Aerodynamics and Gas Dynamics | Iben U.,Robert Bosch GmbH | Munz C.-D.,Institute of Aerodynamics and Gas Dynamics
Journal of Thermal Science | Year: 2016

In the present investigation, we demonstrate the capabilities of the discontinuous Galerkin spectral element method for high order accuracy computation of gas dynamics. The internal flow field of a natural gas injector for bivalent combustion engines is investigated under its operating conditions. The simulations of the flow field and the aeroacoustic noise emissions were in a good agreement with the experimental data. We tested several shock-capturing techniques for the discontinuous Galerkin scheme. Based on the validated framework, we analyzed the development of the supersonic jets during different opening procedures of a compressed natural gas injector. The results suggest that a more gradual injector opening decreases the noise emission. © 2016, Science Press, Institute of Engineering Thermophysics, CAS and Springer-Verlag Berlin Heidelberg.

Edelmann C.A.,University of Stuttgart | Edelmann C.A.,Institute of Aerodynamics and Gas Dynamics | Rist U.,University of Stuttgart | Rist U.,Institute of Aerodynamics and Gas Dynamics
AIAA Journal | Year: 2015

The influence of forward-facing steps on the amplification of disturbances in transonic flows is studied numerically. The main goal is to find a function for an additional amplification factor ΔN that can be incorporated into the eN method. In the present study, both two-dimensional direct numerical simulation of disturbances as well as linear stability theory are used to gain detailed information on the flow. Good agreement between N-factor results from direct numerical simulation and linear stability theory is obtained. Subsonic and slightly supersonic results show differences in the flow topology and in the shape of the N-factor curve, but effective ΔN factors are in the same order of magnitude. Copyright © 2015 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Zahn J.,University of Stuttgart | Zahn J.,Institute of Aerodynamics and Gas Dynamics | Rist U.,University of Stuttgart | Rist U.,Institute of Aerodynamics and Gas Dynamics
AIAA Journal | Year: 2016

Two-dimensional direct numerical simulations are used to study the impact of deep gaps on laminar-turbulent transition in compressible boundary-layer flow. For these, the gap depth-to-width ratio is always larger than five. They are located on a flat plate without pressure gradient.Asteady base flow is used with a Mach number of 0.6, freestream temperature of 288 K, and free-stream pressure of 1 bar. Subsequently, Tollmien-Schlichting waves are introduced by suction and blowing at the wall, and their growth over the gap is evaluated byNfactors. The influence of the gap on laminar-turbulent transition is quantified by the differenceδNcompared with theNfactor obtained for a flat plate without gap. A periodic influence of the gap depth on δNis observed. In the direct numerical simulations, acoustic waves enter the gap and form a standing wave due to reflections, similar as occurring in organ pipes. The feedback of the standing wave on the boundary-layer flow above is essential for the observed δN variations. In a second case, the influence of a specific gap placed in front of a forward-facing step is studied as well. Here, a reduction of theN factor and hence a delay of transition, relative to the flow with step alone, are reached due to the presence of the gap. Copyright © 2015 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Cerretelli C.,General Electric | Wuerz W.,University of Stuttgart | Wuerz W.,Institute of Aerodynamics and Gas Dynamics | Gharaibah E.,General Electric
AIAA Journal | Year: 2010

Fluidic oscillators are actuators that are essentially constituted of a flow vane with no moving parts. They are very effective in generating an oscillating velocity field, and because of their robustness and potential to meet most application requirements they have been thoroughly investigated in previous years. In this work fluidic oscillators have been embedded in an airfoil representative of the outboard sections of wind turbine blades, and subsequently tested at full-scale Reynolds numbers 2:0 · 10 6 ≤ Re ≤ 4:8 · 106 in the laminar wind tunnel at the University of Stuttgart. The effects of the unsteady actuation on the lift and drag strongly depend upon Re, the level of actuation, and the state of the airfoil surface. However, strong improvements have been obtained throughout the whole testing envelope, with relative lift increase spanning from a minimum of 10 to over 60% and substantial stall margin extension. In addition, employing fluidic oscillators strongly reduces the suction surface boundary-layer thickness and the unsteadiness of the mean flow velocity. Copyright © 2010.

Plogmann B.,Institute of Aerodynamics and Gas Dynamics | Wurz W.,Institute of Aerodynamics and Gas Dynamics | Kramer E.,Institute of Aerodynamics and Gas Dynamics
Experiments in Fluids | Year: 2015

The effect of an isolated, cylindrical roughness on the stability of an airfoil boundary layer has been studied based on particle image velocimetry and hot-wire anemometry. The investigated roughness elements range from a sub-critical to a super-critical behavior with regard to the critical roughness Reynolds number. For the sub-critical case, the nonlinear disturbance growth in the near wake is governed by oblique Tollmien–Schlichting (TS) type modes. Further downstream, these disturbance modes are, however, damped with the mean flow stabilization and no dominant modes persist in the far wake. By contrast, in the transitional configuration the disturbance growth is increased, but still associated with a TS-type instability in the near-wake centerline region of the low-aspect (height-to-diameter) ratio element. That is, the disturbances in the centerline region show a similar behavior as known for 2D elements, whereas in the outer spanwise domain a Kelvin–Helmholtz (KH) type, shear-layer instability is found, as previously reported for larger aspect ratio isolated elements. With increasing height and, thereby, aspect ratio of the roughness, the KH-type instability domain extends toward the centerline and, accordingly, the TS-type instability domain decreases. For high super-critical cases, transition is already triggered in the wall-normal and spanwise shear layers upstream and around the roughness. In the immediate wake, periodic shear-layer disturbances roll up into a—for isolated elements characteristic—shedding of vortices, which was not present at the lower roughness Reynolds number cases due to the decreased aspect ratio and, thereby, different instability mechanism. © 2015, Springer-Verlag Berlin Heidelberg.

Plogmann B.,Institute of Aerodynamics and Gas Dynamics | Herrig A.,Institute of Aerodynamics and Gas Dynamics | Wurz W.,Institute of Aerodynamics and Gas Dynamics
Experiments in Fluids | Year: 2013

Discrete frequency tones in the trailing edge noise spectra of NACA 0012 airfoils are investigated with the Coherent Particle Velocity method. The Reynolds number and angle of attack range, in which these discrete frequency tones are present, are consistent with published results. The discrete tones are composed of a main tone and a set of regularly spaced side peaks resulting in a ladder-type structure for the dependency on the free stream velocity. The occurrence of this discrete frequency noise could be attributed to the presence of a laminar boundary layer on the pressure side opening up into a separation bubble near the trailing edge, which was visualized using oil flow. Wall pressure measurements close to the trailing edge revealed a strong spanwise and streamwise coherence of the flow structures inside this laminar separation bubble. The laminar vortex shedding frequencies inferred from the streamwise velocity fluctuations, which were evaluated from hot-wire measurements at the trailing edge, were seen to coincide with the discrete tone frequencies observed in the trailing edge noise spectra. Previous findings on discrete frequency tones for airfoils with laminar boundary layers up to the trailing edge hint at the existence of a global feedback loop. Hence, sound waves generated at the trailing edge feed back into the laminar boundary layer upstream by receptivity and are, then, convectively amplified downstream. The most dominant amplification of these disturbance modes is observed inside the laminar separation bubble. Therefore, the frequencies of the most pronounced tones in the trailing edge noise spectra are in the frequency range of the convectively most amplified disturbance modes. Modifying the receptivity behavior of the laminar boundary layer on the pressure side by means of very thin, two-dimensional roughness elements considerably changes the discrete tone frequencies. For roughness elements placed closer to the trailing edge, the main tone frequency was seen to decrease, while the frequency spacing in-between two successive tones increased. Based on the stability characteristics of the laminar boundary layer and the characteristics of the upstream traveling sound wave, a method for predicting the discrete tone frequencies was developed showing good agreement with the measured results. Hence, with a controlled modification of the laminar boundary layer receptivity behavior, the existence of the proposed feedback loop could be confirmed. At the same time, no significant influence of a second feedback loop previously proposed for the suction side of the NACA 0012 airfoil was observed neither by influencing the boundary layer with a receptivity-roughness element nor by tripping the boundary layer at the leading edge. © 2013 Springer-Verlag Berlin Heidelberg.

Flad D.G.,Institute of Aerodynamics and Gas Dynamics | Frank H.M.,Institute of Aerodynamics and Gas Dynamics | Beck A.D.,Institute of Aerodynamics and Gas Dynamics | Munz C.-D.,Institute of Aerodynamics and Gas Dynamics
20th AIAA/CEAS Aeroacoustics Conference | Year: 2014

The discontinuous Galerkin Spectral Element Method (DG-SEM) is highly attractive for both DNS and LES of turbulent flows due to its low dispersion and dissipation errors, but also because of its good parallel scaling property. We show that especially for underresolved simulations the method has highly beneficial properties for LES, as well as for the direct treatment of the acoustic propagation. We also discuss different approaches for non-reflecting boundary conditions for DG-SEM and show the behavior of the methods for airfoil flows. Our main intend is to directly simulate trailing edge noise for airfoil flows at medium Reynolds numbers.

Filimon A.,Institute of Aerodynamics and Gas Dynamics | Munz C.-D.,Institute of Aerodynamics and Gas Dynamics
Notes on Numerical Fluid Mechanics and Multidisciplinary Design | Year: 2013

In this article we apply the procedure of the iterated defect correction method to the Euler equations as well as to the Navier-Stokes equations. One building block in the defect correction approach is the lower order basic method, usually first or second order accurate. This scheme gives a steady solution of low accuracy as the starting point. The second building block is the WENO reconstruction step to estimate the local defect. The local defect is put into the original equation as source on the right hand side with a minus sign. The resulting modified equation is then again solved with the low order scheme. Due to the source term with the local defect the order of accuracy is iteratively shifted to the order of the reconstruction. We show numerical results for several validation test cases and applications. © 2013 Springer-Verlag Berlin Heidelberg.

Frank H.M.,Institute of Aerodynamics and Gas Dynamics | Munz C.-D.,Institute of Aerodynamics and Gas Dynamics
22nd AIAA/CEAS Aeroacoustics Conference, 2016 | Year: 2016

In low Reynolds number flows such as the flow around a side-view mirror, instability mechanisms in the laminar boundary layer, in particular in laminar separation regions can lead to tonal noise generation. A high order discontinuous Galerkin method is employed to investigate the source mechanisms by means of direct noise computation. In preceding simulations, the flow tones generated by a realistic side-view mirror were reproduced and the underlying mechanism was identified as an aeroacoustic feedback loop. In this paper, we analyze this mechanism by means of simulations at a simplified two-dimensional geometry. The flow field and the acoustic radiation display close similarity to the original mirror case. Furthermore, a global stability analysis provides evidence for the feedback mechanism. For the first time, the typical ladder-structure known from experiments is obtained in numerical simulations. We demonstrate that the tonal modes exhibit close similarities to experimental data in terms of their spectral structure and development with freestream velocity. © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

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