Conference Paper

Experimental and Computational Investigations of Shock Wave – Boundary Layer Interaction on a Transonic Airfoil with and without Bump

Conference Paper

Experimental and Computational Investigations of Shock Wave – Boundary Layer Interaction on a Transonic Airfoil with and without Bump

If you want to read the PDF, try requesting it from the authors.

No full-text available

Request Full-text Paper PDF

To read the full-text of this research,
you can request a copy directly from the authors.

... A similar AC-DBD setup [4] was used and adapted. It will be used at a blowdown type [146] selected parameters with the help of the earlier experiment results are summarized in the A more detailed sketch of the AC-DBD is given in Figure 6.3. The designed DBD actuator will be flush mounted on the plate via spray-on glue, which will help remounting in the future tests. ...
... Layout of the Schlieren Flow Visualization Setup[146] ...
Thesis
Full-text available
Plasma actuators have represented an attractive field for decades and the recent advances in ns-DBDs are now promising new ways to control both the subsonic and supersonic flows with extraordinary efficiency and flexibility. A design methodology, however, doesn’t exist which remains the chief concern of all the researches in the literature. There are many experiments that dealt with changing voltages, frequencies, pulse durations, different voltage waveforms, discrete geometries for the actuators. Even though there is an extensive number of papers suggesting that the dominant mechanism of the actuator is rapid heating, the parameters to maximize the control over the flow, including 2-D interactions, external forces and conditions of operating medium, are not yet well understood and a novel approach is necessary that will incorporate the theory of plasma to understand, model and simulate its physics and experimental results. An extensive literature review was conducted from the theory of MHD to the chemistry of rapid gas heating to unearth the key problems and lacking knowledge about this technology and a basic DBD that will be used in ITU TRL 15 x 15 cm Trisonic Wind Tunnel with Blocks (M4)” nozzle/module was designed to be operative in supersonic conditions and replicate the results of more previous experiments in modifying the boundary layer over a flat plate. At long last, an innovative technique for developing and testing new configurations and necessary experiments for ultimately creating a full methodology for designing a DBD for supersonic flows for different purposes was discussed.
Article
Turbulent flow over a series of increasingly high, two-dimensional bumps is studied by well-resolved large-eddy simulation. The mean flow and Reynolds stresses for the lowest bump are in good agreement with experimental data. The flow encounters a favourable pressure gradient over the windward side of the bump, but does not relaminarize, as is evident from near-wall fluctuations. A patch of high turbulent kinetic energy forms in the lee of the bump and extends into the wake. It originates near the surface, before flow separation, and has a significant influence on flow development. The highest bumps create a small separation bubble, whereas flow over the lowest bump does not separate. The log law is absent over the entire bump, evidencing strong disequilibrium. This dataset was created for data-driven modelling. An optimization method is used to extract fields of variables that are used in turbulence closure models. From this, it is shown how these models fail to correctly predict the behaviour of these variables near to the surface. The discrepancies extend further away from the wall in the adverse pressure gradient and recovery regions than in the favourable pressure gradient region.
Conference Paper
Numerical simulations of internal, transonic flow has been carried out using Large Eddy Simulation. The motion of the unsteady shock and separation has been studied by using Large Eddy Simulations for steady and unsteady boundary conditions. The shock position is highly sensitive to small changes in boundary conditions in the transonic flow range. The shock position and the extent of the separated flow behind it exhibit hysteretic behavior. For the steady-state inflow conditions one observes the presence of certain modes. One of these can be related to the acoustics of the channel, where as another can be related to the shear-layer instability associated with the separation bubble. Copyright © 2008 by the American Institute of Aeronautics and Astronautics, Inc.
Conference Paper
Previous research on the behavior of shock control bumps (SCBs) on transonic airfoils has been largely limited to numerical studies, with experimental investigations primarily limited to basic flow fields in small wind tunnels. This paper examines the possibility of simulating the conditions on a wing in a blow-down supersonic wind tunnel to allow a relatively inexpensive and simple experimental study of the fundamental physics of SCBs. The main requirements are a post-shock adverse pressure gradient and a representative incoming turbulent boundary layer. Tests were carried out at a Mach number of 1.3 using a variety of measurement techniques and the results compared with computations. The ow conditions in the proposed wind tunnel set-up were highly comparable with the computational results for a representative ight condition on a typical transonic airfoil. A contour SCB was tested in the new wind tunnel set-up, and its ow features are discussed. It was found that the SCB brought about an improved total pressure recovery in the boundary layer by the end of the diffuser (corresponding to the airfoil trailing edge) and this was attributed to a vortical wake generated by baroclinic effects. This provides direct evidence in support of the suggestion that SCBs could also be used as a form of boundary layer control.
Article
Because of their ubiquitous presence in high-speed flight and their impact on vehicle and component performance, shock-wave/boundary-layer interactions have been studied for about 50 years. Despite truly remarkable progress in computational and measurement capabilities, there are still important quantities that cannot be predicted very accurately, that is, peak heating in strong interactions, or cannot be predicted at all, that is, unsteady pressure loads. There remain observations that cannot be satisfactorily explained and physical processes that are not well understood. Much work remains to be done. Based on the author's own views and those of colleagues, some suggestions are made as to where future efforts might be focused.,lust as the first workers in the field could not have foreseen the capabilities generated by the computer/instrumentation revolution of the past 20 years, it is probably fair to assume that the extent of our vision and imagination in the year 2000 is equally limited. New simulation and measurement techniques will doubtlessly become available in the next 10 or 20 years, the results from which will render many of the current concerns moot. However, as vehicle missions and cost constraints become ever more demanding, How regimes harsher, and flow control/manipulation becomes an absolute necessity, the need for an ever deeper physical understanding and a more accurate, more robust simulation capability will only grow. Now is the time to lay the groundwork for the next 50 years.
Article
This aim of this study was to investigate the phenomenon of shock/boundary-layer interaction on a supercritical airfoil with and without shock-induced separation and to compare the results obtained by a conventional boundary-layer probe with those from a two-component Laser-Doppler Anemometer. Measurements were made in the 1 multiplied by 1 m**2 transonic wind tunnel (TWG) of the DFVLR Goettingen on a 250-mm-chord model of the airfoil CAST 7/DOA1 at M// infinity equals 0. 765 and angles of attack of alpha equals 3. 6 and 5 deg. The Reynolds number was Re equals 2. 5 multiplied by 10**6. Generally, the agreement between the two data sets was satisfactory. However, some discrepancies remained, especially at positions close to the wall, where the LDA data showed, in some instances, a considerable scatter. Furthermore, within the separated region where flow reversal occurs, both the LDA and the probe data are suspect.
Article
Results are presented of four research programs concerned with the phenomenon of shock boundary-layer interaction and ways that such interactions affect scaling of transonic airfoil flows. The results were obtained using a diverging nozzle and several simulated airfoil contours. Main variables of these programs were the freestream Mach number, the Reynolds number, the mode of boundary-layer transition, and the geometry and location of boundarylayer tripping devices. Emphasis in presenting the results is placed on the flow development that may lead to large transonic scale effects and on the dependency of this development on the unit Reynolds number and the initial boundary-layer condition. An attempt is made to relate components of the shock boundary-layer interaction, i.e., the pressure rise to shockinduced separation, the length of the separation bubble, and the extent of rear separation to parameters associated with the boundary-layer condition upstream of the shock. © 1971, American Institute of Aeronautics and Astronautics, Inc., All rights reserved. © 1971, American Institute of Aeronautics and Astronautics, Inc., All rights reserved.
Article
There are few successful computational reports for transonic airfoil now worked out with the pressure-based method. In this study, an advanced approach based on a pressure correction scheme is developed to solve the Reynolds-averaged Navier-Stokes equations for turbulent transonic now around the airfoil RAE 2822. An implicit numerical dissipation model is adopted to create a dissipation mechanism based on pressure gradients to damp the destabilizing numerical effects, without smearing the physical discontinuity at shocks. The standard k-epsilon turbulence closure with a near-wall one-equation model is used. The computational results are compared with experimental data. Several discretization schemes such as the second-order upwind, hybrid, and MUSCL schemes for convection terms are investigated. The computational results show that the proposed pressure-based method has a resolution comparable to, or better than, the traditional time-marching methods.
Article
Large-eddy simulation (LES) has been used to calculate the flow of a statistically two-dimensional turbulent boundary layer over a bump. Subgrid-scale stresses in the filtered Navier Stokes equations were closed using the dynamic eddy viscosity model. LES predictions for a range of grid resolutions were compared to the experimental measurements of Webster et al. (1996). Predictions of the mean flow and turbulence intensities are in good agreement with measurements. The resolved turbulent shear stress is in reasonable agreement with data, though the peak is over-predicted near the trailing edge of the bump. Analysis of the flow confirms the existence of internal layers over the bump surface upstream of the summit and along the downstream trailing at plate, and demonstrates that the quasi-step increases in skin friction due to perturbations in pressure gradient and surface curvature selectively enhance near-wall shear production of turbulent stresses and are responsible for the formation of the internal layers. Though the flow experiences a strong adverse pressure gradient along the rear surface, the boundary layer is unique in that intermittent detachment occurring near the wall is not followed by mean-flow separation. Certain turbulence characteristics in this region are similar to those previously reported in instantaneously separating boundary layers. The present investigation also explains the driving mechanism for the surprisingly rapid return to equilibrium over the trailing flat plate found in the measurements of Webster et al. (1996), i.e. the simultaneous uninterrupted development of an inner energy-equilibrium region and the monotonic decay of elevated turbulence shear production away from the wall. LES results were also used to examine response of the dynamic eddy viscosity model. While subgrid-scale dissipation decreases/increases as the turbulence is attenuated/enhanced, the ratio of the (averaged) forward to reverse energy transfers predicted by the model is roughly constant over a significant part of the layer. Model predictions of backscatter, calculated as the percentage of points where the model coefficient is negative, show a rapid recovery downstream similar to the mean-flow and turbulence quantities.
Article
Benchmark experimental data obtained in the two-dimensional, transonic flow field surrounding a supercritical airfoil are presented. Airfoil surface and tunnel wall pressure and LDV measurements are used to describe the flow on the model, above the wing and in the wake. Comparisons are made with calculations using the Reynolds-averaged Navier-Stokes equations. The results illustrate the performance of two turbulence models in both separated and attached flows. The largest differences between theory and experiment occurred in separated flows with the Johnson and King turbulence model providing the best estimates.
Article
Numerical simulations of transonic airfoil flows using the Reynolds-averaged Navier-Stokes equations and various turbulence models are presented and compared with experimental data. Three different airfoils were investigated under varying flow conditions ranging from subcritical unseparated flows to supercritical separated flows. The turbulence models investigated consisted of three zero-equation models and one two-equation model. For unseparated flows involving weak viscous-inviscid interactions, the four models were comparable in their agreement with experiment. For separated flows involving strong viscous-inviscid interactions, the nonequilibrium zero-equation model of Johnson and King gave the best overall agreement with experiment.
Advisory Report No. 138 EXPERIMENTAL DATA BASE FOR COMPUTER PROGRAM ASSESSMENT
  • Agard
AGARD, "Advisory Report No. 138 EXPERIMENTAL DATA BASE FOR COMPUTER PROGRAM ASSESSMENT," Technical Editing and Reproduction Ltd., London, 1979.
Numerical and Experimental Validation of Three-Dimensional Shock Control Bumps
  • B König
  • M Pätzold
  • T Lutz
  • E Krämer
  • E Rosemann
  • K Richter
  • H Uhlemann
B. König, M. Pätzold, T. Lutz, E. Krämer, E. Rosemann, K. Richter and H. Uhlemann, "Numerical and Experimental Validation of Three-Dimensional Shock Control Bumps," in 4th Flow Control Conference, Seattle, Washington, 2008.