Conference Paper

Vortical structures on three-dimensional shock control bumps

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Abstract

The vortical wake structure produced by a three-dimensional shock control bump (SCB) is thought to be useful for controlling transonic buffet on airfoils. However, at present the vorticity produced is relatively weak and the production mechanism is not well understood. Using a combined experimental and computational approach, a preliminary investigation on the wake vorticity for different bump geometries has been carried out. The structure of the wake for on and off-design conditions are considered, and the effects on the downstream boundary layer demonstrated. Three main vortical structures are observed: a primary vortex pair, weak inter-bump vortices and shear flow in the lambda-shock region. The effect of pressure gradients on vortex strength is examined and it is found that spanwise pressure gradients on the front section of the bump are the most significant parameter influencing vortex strength.

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... However, the real situation is rather more complex, due primarily to the fact that 3-D SCBs have been observed to introduce streamwise vorticity into the flow [17,26,38]. The precise mechanism by which this vorticity is generated is still a subject of research, although a recent publication suggesting that the spanwise pressure gradients present on a 3-D SCB are of key importance [9] is perhaps the most convincing to date. The flow physics of 3-D SCBs can become significantly more complex when the shock moves away from its optimal location. ...
... It is therefore unsurprising that much of the experimental work that has been performed on SCBs has been done using bumps mounted on the floor of small-scale supersonic wind tunnels-notably [6,8,9,12,26,38]. The advantage of this is that the tests are inexpensive, repeatable and concentrate solely on the flow local to the bump, thus enabling a detailed examination. ...
... A prominent flow feature associated with 3-D SCBs is the vortical wake structure. Whilst the most striking vortex formation is caused by flow separation when the SCB is operating under off-design conditions, the bumps also produce vortices when operating under design conditions which are not only observable, but also produce measurable effects on the downstream flow [6,8,9,24,29,38]. In its guise as a drag reduction device, it can be argued that the presence of streamwise vortices is not welcome, since it will add a (small) element of drag. ...
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This paper presents a review of research on shock control bumps (SCBs), a class of flow control device with potential for application to transonic wings. Beginning with a brief review of the origins of the SCB concept, the primary focus is on the more recent studies from the last decade. Results from both experimental and numerical work are considered and the synergy between these two approaches to SCB research is critically explored. It is shown that the aerodynamic performance enhancement potential of SCBs, namely their capacity for drag reduction and delaying the onset of buffet for transonic wings, has been widely demonstrated in the literature, as has the high sensitivity of SCB performance to flow conditions including shock strength and position, and post-shock adverse pressure gradient. These characteristic features of SCBs are relatively well explained in terms of the flow physics that have been observed for different bump geometries. This stems from a number of studies that have focused on the balance of viscous and inviscid flow features and also the mechanism by which finite span SCBs generate streamwise vorticity. It is concluded that our understanding of SCBs is reaching an advanced level of maturity for SCBs in simple configurations and steady flow fields. However, SCB performance in unsteady flow and on swept wings requires further investigation before the concept can be considered a viable candidate for transonic wings. These investigations should adopt a multi-disciplinary approach combining carefully designed experiments and targeted computations. Finally, two concepts for future SCB research are suggested: the adaptive SCB and SCBs in engine intakes.
... One option is to use an array of finite span (3-D) SCBs, which have been shown to be more robust to variations in shock position while still achieving an on-design performance benefit close to that of an optimal 2-D SCB [10]. Furthermore, some studies have suggested that 3-D SCBs may also delay the onset of transonic buffet (relative to a clean wing or one with a 2-D SCB), and thus offer potential for enhancing off-design wing performance [6,11]. ...
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Chapter
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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 size specifications for suitable tracer particles for particle image velocimetry (PIV), particularly with respect to their flow tracking capability, are discussed and quantified for several examples. A review of a wide variety of tracer materials used in recent PIV experiments in liquids and gases indicates that appropriately sized particles have normally been used. With emphasis on gas flows, methods of generating seeding particles and for introducing the particles into the flow are described and their advantages are discussed.
Conference Paper
This volume contains results of the German CFD initiative MEGAFLOW which combines many of the CFD development activities from DLR, universities and aircraft industry. It high-lights recent improvements and enhancement of the MEGAFLOW software system. This software system includes the block-structured Navier-Stokes code FLOWer and the unstructured Navier-Stokes code TAU. Besides improvments to numerical algorithms and physical modelling capabilities of these codes the validation of their capability to predict viscous flows around complex industrial applications for transport aircraft design is presented. This documents the high level of maturity both codes have reached. The goal of MEGAFLOW to develop and validate dependable and efficient numerical tools for the aerodynamic simulation of complete aircraft is proven based on the intensive use of FLOWer and TAU by the German aerospace industry in the design process of a new aircraft.
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
Three-dimensional bumps have been developed and investigated on transonic wings, aiming to fulfill two major objectives of shock-wave/boundary-layer interaction control, that is, drag reduction and buffet delay. An experimental investigation has been conducted for a rounded bump in channel flow at the University of Cambridge and a computational study has been performed for a spanwise series of rounded bumps mounted on a transonic aerofoil at the University of Stuttgart. In both cases wave drag reduction and mild control effects on the boundary layer have been observed. Control effectiveness has been assessed for various bump configurations. A double configuration of narrow rounded bumps has been found to perform best, considerably reducing wave drag by means of a well-established λ-shock structure with little viscous penalty and thus achieving a maximum overall drag reduction of about 30 %, especially when significant wave drag is present Counter-rotating stream wise vortex pairs have been produced by some configurations as a result of local flow separation. On the whole a large potential of three-dimensional control with discrete rounded bumps has been demonstrated both experimentally and numerically.
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.