Chapter

Numerical and Experimental Examination of Shock Control Bump Flow Physics

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Abstract

A method allowing a detailed investigation of the flow physics of shock control bumps (SCBs) on an unswept airfoil has been developed by comparison of the results of experiments and computations. A simple wind tunnel set-up is proposed which is shown to generate representative baseline conditions, allowing fine details of the flow to be measured using an array of techniques. Computational data for the same bump configuration is then validated against the experimental results, allowing a more intimate analysis of the flow physics as well as relating wind tunnel results to the performance of the SCB on an unswept wing.

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Article
In addition to efficient reduction of wave drag in transonic flight, shock control bumps also offer some potential for buffet alleviation. In the present paper, two different approaches for buffet control by shock control bumps are compared and assessed based on time-resolved (unsteady) Reynolds-averaged Navier–Stokes simulations: downstream positioning of two-dimensional bumps, necessitating an adaptive device for combining both features (wave drag reduction and buffet control), and vortex generation by three-dimensional bumps as “smart vortex generators”. With the intention of providing detailed insight into dominant flow features and linking geometrical bump characteristics to those flow structures, the effect of bump design condition, crest height, and streamwise positioning of two- and three-dimensional shock control bumps on buffet behavior and performance of a supercritical, unswept wing section has been analyzed. Two-dimensional shock control bumps improve buffet behavior by an efficient shock strength reduction in combination with positive effects on flow separation. For three-dimensional bumps, the same buffet-affecting mechanisms have been observed, only less dominant due to the finite spanwise extent. Furthermore, it has been demonstrated that the strength of three-dimensional bumps’ vortical wake can be tuned by appropriate bump shaping and that this strength positively correlates with delayed buffet onset.
Chapter
Originally developed as a flow control device Shock Control Bumps (SCB) reduce wave drag of an aircraft wing at off-design in transonic speed effectively. Recently, another field of application for such bumps has been studied, namely the delay and alleviation of buffet, an unsteady shock motion due to continuous flow separation and re-attachment at the rear part of the airfoil. In principle the idea of buffet alleviation is the use of SCB as a sort of ‘smart’ vortex generator. Considerable effort has been undertaken to link geometrical bump features to buffet affecting flow characteristics. In this paper a parametric study on the influence of flank shape of a three-dimensional wedge-shaped SCB on its performance and buffet behavior is presented. It has been found that performance as well as buffet behavior can be improved by optimization of the bump flanks. The study shows that length of front and rear flank should be increased up to given constraints (e.g. flaps on a wing or inserts for a wind tunnel model) and a narrow front and wide rear flank increase c L, max and damp lift oscillations at buffet onset.
Article
Full-text available
Three-dimensional shock control bumps have long been investigated for their promising wave drag reduction capability. However, a recently emerging application has been their deployment as "smart" vortex generators, which offset the parasitic drag of their vortices against their wave drag reduction. It is known that three-dimensional shock control bumps produce streamwise vortices under most operating conditions; however, there have been very few investigations that have aimed to specifically examine the relevant flow structures. In particular, the strength of the vortices produced as well as the factors influencing their production are not well known. This paper uses a joint experimental and computational approach to test three different shock control bump shapes, categorizing their flow structures. Four common key vortical structures are observed, predominantly shear flows, although all bumps also produce a streamwise vortex pair. Both cases with and without flow separation on the bump tails are scrutinized. Finally, correlations between the strength of the main wake vortices and pressure gradients at various locations on the bumps are assessed to investigate which parts of the flow control the vortex formation. Spanwise flows on the bump ramp are seen to be the most influential factor in vortex strength. © Copyright 2015 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
Chapter
There are a number of engineering optimization problems that are single-objective or can be represented by a weighted sum if the weights are known in advance. Inverse design including reconstruction problems are also another type of single-objective problems. This chapter illustrates the application of some of the techniques described in previous chapters to wing airfoil reconstruction design, active flow control bump design, and generic aircraft wing airfoil section design optimizations.
Article
Full-text available
A method of accurately calculating transonic and low Reynolds number airfoil flows, implemented in the viscous-inviscid design/analysis code ISES, is presented. The Euler equations are discretized on a conservative streamline grid and are strongly coupled to a two-equation integral boundary-layer formulation, using the displacement thickness concept. The entire discrete equation set, including the viscous and transition formulations, is solved as a fully coupled nonlinear system by a global Newton method. This is a rapid and reliable method for dealing with strong viscous-inviscid interactions, which invariably occur in transonic and low Reynolds number airfoil flows. The results presented demonstrate the ability of the ISES code to predict transitioning separation bubbles and their associated losses. The rapid airfoil performance degradation with decreasing Reynolds number is thus accurately predicted.
Conference Paper
Full-text available
Within the European Project Telfona the Pathfinder Model was designed, analyzed numerically, constructed and tested with the aim of obtaining a laminar flow testing capability in the European Transonic Wind Tunnel (ETW). The model was designed for natural laminar flow (NLF) for transonic flow conditions with high Reynolds number. Results of pre-test numerical analysis demonstrated that the Pathfinder wing pressure distribution was adequate for providing calibration test points. The ETW tests provided pressure distribution data while transition positions were determined from images using the Cryogenic Temperature Sensitive Paint Method (cryoTSP). The evaluation of this data with several transition prediction tools was used to establish the transition N-factor values for ETW. In this work, after-test CFD solutions are obtained using numerical Navier-Stokes solutions. In the first part of this work, numerical results are given which verify the requirements of the Pathfinder wing as a calibration model. In the second part, it is shown that for selected flow conditions a good agreement is obtained between stability analysis based on experimental and numerical data. In the third part the correlation of experimental transition locations to critical N-factors is summarized for ETW Test Phases I and II. In the fourth part numerical analysis and experimental data are used complementarily.
Article
Full-text available
Numerical and experimental studies have been performed to show the potential for drag reductions of an array of discrete three-dimensional shock control bumps. The bump contour investigated was specifically designed by means of CFD-based numerical optimization for wind tunnel testing on a modern transonic airfoil. The experimental investigations focused on turbulent flow at a Reynolds number of 5 million and were carried out at the Transonic Wind Tunnel G¨ottingen. Drag reductions of around 10% in the drag-rise region were found in the experiment even though the results were influenced by wind tunnel interference effects. A detailed numerical study of the wind tunnel environment reproduced the influence of the wind tunnel walls on the bump performance and gave good agreement to the experimental results.
Book
This volume contains a thorough description of the EU-supported project EUROSHOCK II concerned with the investigation of active shock and boundary layer control to improve aircraft performance. Discussed are basic experiments, supplemented by Navier-Stokes computations, to improve and validate physical models relevant to control and the extension, validation and application of various computational methods to airfoil and wing flows with control. Furthermore described are experiments on airfoils and wings carried out to assess the aerodynamic benefits of control and to provide data for validation purposes. Finally, control applications to real aircraft and the corresponding installation penalties and mission benefits are addressed. This volume is a sequel to Vol. 56 on passive shock control.
Article
This paper reviews and highlights recent developments of certain aspects of flow control concerned with reducing the drag of, and delaying flow separation on, wings and bodies over which the flow is turbulent. The study is restricted to devices that extend beyond the viscous sub-layer but are on a smaller scale than geometric features of the aircraft (e.g. wing chord). The review is mainly concerned with developments within the UK, although significant developments in other countries are discussed. The review discusses types of flow that need to be controlled, basic features of flow control devices and applications. It concludes with recommendations for future research.
Conference Paper
This paper describes a fundamental experimental study of the flow structure around a single three-dimensional (3D) transonic shock control bump (SCB) mounted on a flat surface in a wind tunnel. Tests have been carried out with a Mach 1.3 normal shock wave located at a number of streamwise positions relative to the SCB. Details of the flow have been studied using the experimental techniques of schlieren photography, surface oil flow visualization, pressure sensitive paint, and laser Doppler anemometry. The results of the work build on the findings of previous researchers and shed new light on the flow physics of 3D SCBs. It is found that spanwise pressure gradients across the SCB ramp and the shape of the SCB sides affect the magnitude and uniformity of flow turning generated by the bump, which can impact on the spanwise propagation of the quasi-two-dimensional (2D) shock structure produced by a 3D SCB. At the bump crest, vortices can form if the pressure on the crest is significantly lower than at either side of the bump. The trajectories of these vortices, which are relatively weak, are strongly influenced by any spanwise pressure gradients across the bump tail. A significant difference between 2D and 3D SCBs highlighted by the study is the impact of spanwise pressure gradients on 3D SCB performance. The magnitude of these spanwise pressure gradients is determined largely by SCB geometry and shock position.
Conference Paper
The desire to design the most efficient transport aircraft has lead to many different attempts to minimise drag. One approach is the use of shock control bumps (SCBs), which reduce the drag associated with local shockwaves on transonic wings. In particular, three-dimensional (3D) SCBs have gained popularity as simple, efficient and robust devices capable of reducing the wave drag over transonic wings. This paper presents a computational study of the performance of 3D SCBs on a transonic wing, relating key SCB design variables to the overall wing aerodynamic performance. An efficient parameterisation scheme allows 3D SCBs to be directly compared to 2D designs, indicating that if compared solely by on-design aerodynamic performance, 2D SCBs form the limiting case of all 3D designs. An off-design study indicates that again, 2D SCBs are capable of improved performance over comparable 3D designs. The major advantage of 3D SCBs lies in the production of streamwise vortices, which could be beneficial in terms of control of local separation under off-design conditions. Vortex production by 3D SCBs is identified and key SCB geometric parameters are shown to be able to control the strength of these vortices. The results indicate that the true benefit of 3D SCBs is likely to be as an ‘efficient vortex generator’, combining on-design drag reduction with off-design separation / buffet control.
Conference Paper
Robustness enhancement for Shock Control Bumps (SCBs) on transonic wings is an ongoing topic because most designs provide drag savings only in a relatively small band of the airfoil polar. In this paper, different bump shapes are examined with CFD methods which are validated first by comparison with wind tunnel results. An evaluation method is introduced allowing the robustness assessment of a certain design with little computational effort. Shape optimizations are performed to trim SCB designs to maximum performance on the one hand and maximum robustness on the other hand. The results are analysed and different and parameters influencing the robustness are suggested.
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
The aerospace industry increasingly relies on advanced numerical simulation tools in the early design phase. This volume provides the results of a German initiative which combines many of the CFD development activities from the German Aerospace Center (DLR), universities, and aircraft industry. Numerical algorithms for structured and hybrid Navier-Stokes solvers are presented in detail. The capabilities of the software for complex industrial applications are demonstrated.
Article
A vortex-skeleton (VS) modeling technique for three-dimensional separated flows is developed and tested by means of the electromagnetic analogy. The VS method applies the Biot-Savart law to approximate the surface streamline patterns and flow topologies determined analytically using streamsurface bifurcations by Hornung and Perry (1984). Despite the simplifications involved, VSs are shown to reproduce the essential features observed in separated flows. The VS method is tested experimentally using the analogy between the velocity field around a potential line vortex and the magnetic field around a wire carrying a current (both of which obey the Biot-Savart law). The results are presented graphically in wiring diagrams, VS diagrams, and photographs of the magnetic-field lines. The flow configurations observed are classified, and four owl-face patterns are discussed in detail.
Article
A detailed study of transonic flow over three-dimensional bumps has been conducted using experimental measurements and computational simulation. The aim of the investigation is to determine the flow characteristics over these potentially effective flow control devices for wave drag reduction for transonic aircraft wings. Careful qualitative and quantitative matching of the simulation and experimental conditions has considerably improved the agreement between them. In both the experiments and the computation, the shock position in the working section was found to be sensitive not only to the back pressure and the sidewall effects, but also to the incoming boundary layer characteristics. Two turbulence models, a modified Baldwin–Lomax model with enhanced performance in separated flow (curvature model) and a two-equation model (k–ω) have been implemented in the numerical simulation. Overall, both turbulence models gave reasonable results for the uncontrolled and controlled cases regarding the inviscid flowfield and the shock structure. In particular, a pair of streamwise vortices embedded in the boundary layer was captured by the two models. The traces in the surface oil flow pictures from the experiment also suggested the existence of the streamwise vortices. The combined surface and flowfield data provide some further insight into the flow physics on the shock control ramp bump, which is discussed in the paper. In addition, the study also demonstrates the enhanced capability of the algebraic model in capturing separated flow features with lower computational cost as compared to the k–ω model.
Article
A novel supersonic wind tunnel setup is proposed to enable the investigation of control on a normal shock wave. Previous experimental arrangements were found to suffer from shock instability. Wind tunnel tests with and without control have confirmed the capability of the new setup to stabilise a shock structure at a target position without changing the nature of the shock wave / boundary layer interaction flow at M∞ = 1.3 and M ∞ = 1.5. Flow visualisation and pressure measurements with the new setup have revealed detailed characteristics of shock wave / boundary layer interactions and a λ-shock structure as well as benefits of control in total drag reduction in the presence of 3D bump control.
Article
The use of wind-tunnel setup for study of normal shock wave/boundary layer interaction control, was investigated. The rectangular working section that was 114 mm wide, and 178 mm high at the straight downstream of the nozzle was used. The incoming airflow was partitioned by a plate of 6 mm thickness to overcome the problem of shock wave instability. The height of the upper and lower passage was maintained at 91 and 122 mm respectively. The incoming boundary layer thickness was 5.7 mm and the Reynolds number based on boundary-layer displacement thickness was approximately 25,000. It was observed that shock can be located above 3-D bump and large λ-shock structure whose front shock leg starts at the onset of control cab be analyzed. Result shows that wind-tunnel setup can be used to test various types of shock control at positions where conventional setups are unable to hold shock system due to shock instability.
Article
The law of the wake and the law of the wall in incompressible turbulent boundary layers formulated by Coles (1956) and their use by Mathews et al. (1970) in the development of a wall-wake representation of the velocity profile in a form applicable for isoenergetic compressible boundary layers are extended to a modified wall-wake velocity profile for turbulent compressible boundary layers. The modified wall-wake profile is shown to provide good representations of experimental velocity distributions.
Article
Two new two-equation eddy-viscosity turbulence models will be presented. They combine different elements of existing models that are considered superior to their alternatives. The first model, referred to as the baseline (BSL) model, utilizes the original k-omega model of Wilcox In the inner region of the boundary layer and switches to the standard k -epsilon model in the outer region and in free shear flows. It has a performance similar to the Wilcox model, but avoids that model's strong freestream sensitivity. The second model results from a modification to the definition of the eddy-viscosity in the BSL model, which accounts for the effect of the transport of the principal turbulent shear stress. The new model is called the shear-stress transport-model and leads to major improvements in the prediction of adverse pressure gradient flows.
Drag reduction by shock boundary layer control
  • E Stanewsky
  • J Délery
  • J Fulker
  • P De Matteis
Auslegungsstudien von 3-D Shock-Control-Bumps mittels numerischer Optimierung (Design of 3D shock control bumps by numerical optimisation
  • M Pätzold