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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. ...

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]. ...

This paper presents the results from a study to design an adaptive shock control bump for a transonic aerofoil. An optimisation framework comprising aerodynamic and structural computational tools has been used to assess the performance of candidate adaptive bump geometries based on a novel surface-pressure-based performance metric. The geometry of the resultant design is a unique feature of its adaptivity; being strongly influenced by the (passive) aerodynamic pressure forces on the flexible surface as well as the (active) displacement constraints. This optimal geometry bifurcates the shock-wave and carefully manages the recovering post-shock flow to maximise pressure-smearing in the shock-region with only a small penalty in L/D for the aerofoil. Short adaptive bumps (with small imposed displacements) generally perform better than taller ones, and maintain their performance advantage for a wide range of bump positions, suggesting good robustness to variations in shock position, which are an inevitable feature of a real-world flight application. Such devices may offer advantages over conventional (fixed geometry) shock control bumps, where optimal performance is achieved with taller devices, at the expense of poor robustness to variations in shock position.

Interactions between vortices and shock waves, which are often encountered in transonic and supersonic flows, can cause the vortices to break down or burst. Experiments aimed at better understanding these interactions are performed in the range Mach 1.3 to Mach 1.5, with a particular focus on obtaining validation-quality reference data. The resulting measurements provide valuable information to improve prediction of normal shock-induced vortex breakdown at these Mach numbers through a number of physical observations. Firstly, high-speed schlieren visualisation demonstrates that there is a direct link between vortex breakdown and local perturbations in the wave pattern. In addition, a new type of interaction is identified at Mach 1.3, characterised by intermittent breakdown events which are likely caused by feedback between vortex breakdown and adverse pressure gradient. Finally, at the higher Mach numbers, the favourable streamwise pressure gradient imposed by the separation of the tunnel wall boundary layers appears to stabilise the vortex and delay the onset of breakdown.

The quadratic constitutive relation is a simple extension to the linear eddy-viscosity hypothesis and has shown some promise in improving the computation of flow along streamwise corner geometries. In order to further investigate these improvements, the quadratic model is validated by comparing RANS simulations of a Mach 2.5 wind tunnel flow with high-quality experimental velocity data. Careful set up and assessment of computations using detailed characterisation data of the overall flow field suggests a minimum expected discrepancy of approximately 3% for any experimental-computational velocity comparisons. The corner regions of the rectangular cross-section wind tunnel exhibit velocity differences of 7% between experimental data and computations with linear eddy-viscosity models, but these discrepancies are reduced to 4-5% when the quadratic constitutive relation is used. This improvement can be attributed to a better prediction of the corner boundary-layer structure, due to computations reproducing the stress-induced streamwise vortices which are known to exist in this flow field. However, the strength and position of these vortices do not correspond exactly with those in the measured flow. A further observation from this study is the appearance of additional, non-physical vortices when the value of the quadratic coefficient in the relation exceeds the recommended value of 0.3.

Streamwise-coherent structures were observed in schlieren images of a Mach 2.5 flow in an empty supersonic wind tunnel with a rectangular cross section. These features are studied using Reynolds-averaged Navier–Stokes computations in combination with wind-tunnel experiments. The structures are identified as regions of streamwise vorticity embedded in the sidewall boundary layers. These vortices locally perturb the sidewall boundary layers, and they can increase their thickness by as much as 37%. The vortices are caused by a region of separation upstream of the nozzle where there is a sharp geometry change, which is typical in supersonic wind tunnels with interchangeable nozzle blocks. Despite originating in the corners, the vortices are transported by secondary flows in the sidewall boundary layers so they end up near the tunnel center height, well away from any corners. The successful elimination of these sidewall vortices from the flow is achieved by replacing the sharp corner with a more rounded geometry so that the flow here remains attached.

In order to be able to judge the effectiveness of transition induction in WP-2, reference flow cases were planned in WP-1. There are two obvious reference cases—a fully laminar interaction and a fully turbulent interaction. Here it should be explained that the terms “laminar” and “turbulent” interaction refer to the boundary layer state at the beginning of interaction only. There are two basic configurations of shock wave boundary layer interaction and these are a part of the TFAST project. One is the normal shock wave, which typically appears at the transonic wing and on the turbine cascade. The characteristic incipient separation Mach number range is about M = 1.2 in the case of a laminar boundary layer and about M = 1.32 in the case of turbulent boundary layer. The second typical flow case is the oblique shock wave reflection. The most characteristic case in European research is connected to the 6th FP IP HISAC project concerning a supersonic business jet. The design speed of this airplane is M = 1.6. Therefore the TFAST consortium decided to use this Mach number as the basic case. Pressure disturbance at this Mach number is not very high and can be compared to the disturbance of the normal shock at the incipient separation Mach number mentioned earlier. As mentioned earlier, shock reflection at M = 1.6 may be related to incipient separation. Therefore two additional test cases were planned with different Mach numbers. ITAM conducted an M = 1.5 test case, and TUD an M = 1.7 test case. These partners have also previously made very specialized and successful contributions to the UFAST project.

Study of transition location effect (from natural transition to fully turbulent) on separation size, shock structure and unsteadiness was the focus of this WP. Boundary layer tripping (by wire or roughness) and flow control devices (VG) were used for boundary layer transition induction. Although this type of flow field had been studied widely in the past, there remains considerable uncertainty on the effects of transition on transonic aerofoil performance. In particular it is not known how close to the shock location transition has to occur to avoid detrimental effects associated with laminar shock-induced separation. Furthermore, it was unclear how best to provoke transition on an airfoil featuring significant laminar flow and how close to the shock this needs to be performed. Finally, current CFD methods are particularly challenged by such transitional flows. In this work package some of the findings from the basic research performed in other WPs was applied. Specialized large-scale transonic wind tunnels running cost is very high therefore using such facilities is not appropriate for upstream research programs such as TFAST. Therefore we have used existing wind tunnels within our consortium. One of these is a transonic test section at UCAM where laminar and transitional profiles were studied previously at Reynolds numbers up to 2 million (based on chord length). This wind tunnel allowed basic investigations of the transition location effects on a shock induced separation and unsteadiness for a relatively large number of parameters. A larger wind tunnel at Institute of Aviation in Warsaw was used, which enabled the investigation of a much larger aspect ratio profile. In this facility it was possible to measure a whole force polar up to and including the buffet boundary. The research was carried out for the natural b/l transition location as well as different methods of tripping.

The interaction between a normal shock wave and a boundary layer is investigated over a curved surface for a Reynolds number range, based on boundary-layer growing length x, of \(0.44\times 10^6\le \text {Re}_x\le 1.09\times 10^6\). The upstream boundary layer develops around the leading edge of the model before encountering a \(M\) \(\sim \)1.4 normal shock. This is followed by adverse pressure gradients. The shock position and strength are kept constant as \(\text {Re}\) is progressively varied. Infra-red thermography is used to determine the nature of the upstream boundary layer. Across the \(\text {Re}\) range, this is observed to vary from fully laminar to fully turbulent across the entire span. Regardless of the boundary-layer state, the interaction remains benign in nature, without large scale shock-induced separation or unsteadiness. Schlieren images show a pronounced oblique wave developing upstream of the main shock for the laminar cases, this is believed to correspond to the separation and subsequent transition of the laminar shear layer. Downstream of the shock, in the presence of adverse pressure gradients, the boundary-layer growth rate is inversely proportional to \(\text {Re}\). Nonetheless, across the entire range of inflow conditions the boundary layer recovers quickly to a healthy turbulent boundary layer. This suggests the upstream boundary-layer state, and its transition mechanism, to have little effect on the outcome of its interaction with a normal shock wave.
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The flowfield around five transonic inlet lips at high incidence is investigated for a variety of flow conditions around a design point representative of high-incidence maneuvering. Changes to the operating point are simulated by varying the angle of incidence and the mass flow rate over the lip, which is intended to mimic the effect of an increase in engine flow. For these inflow conditions, the flow on the lip is characterized by a supersonic region, which is terminated by a near-normal shock wave. Of particular interest is the effect of the lip geometry and operating point on the boundary layer at the equivalent fan location. The parametric investigation revealed a significant effect of the lip shape on the position and severity of the shock-wave/boundary-layer interaction. From correlation studies, it appears that the extent of shock-induced separation is the main factor affecting the boundary-layer state downstream of the normal shock-wave/boundary-layer interaction. Somewhat surprisingly, this was found to be independent of shock strength.

During high-incidence maneuvers, shock-wave/boundary-layer interactions can develop over transonic inlet lower lips, significantly impacting aerodynamic performance. Here, a novel experimental rig is used to investigate the nature and severity of these interactions for a typical high-incidence scenario. Furthermore, the sensitivity to changes in angle of incidence and mass-flow rate is explored, as potentially experienced across off-design operations. The reference flowfield, informed by typical climb conditions, is defined by an incidence of 23 deg and a freestream Mach number M=0.435. The lower-lip flow is characterized by a rapid acceleration around the leading edge and an M≈1.4 shock ahead of the intake diffuser. Overall, this flowfield is found to be relatively benign, with minimal shock-induced separation. Downstream of the interaction, the boundary layer recovers a healthy profile ahead of the nominal fan location. Increasing the incidence by 2 deg, the separation becomes noticeably larger and unsteadiness develops. Detrimental effects are exacerbated at an even higher incidence of 26 deg. Increasing the mass-flow rate over the lip by up to 15% of the initial value has minor effects on performance and is not found to inhibit the boundary-layer profile recovery.

In a rectangular cross-section wind tunnel, a separated oblique shock reflection is set to interact with the turbulent boundary layer (oblique shock wave/turbulent boundary layer interaction (SBLI)) both on the bottom wall and in the corners formed by the intersection of the floor with the sidewalls. To examine how corner separations can affect the ‘quasi-two-dimensional’ main interaction and by what mechanisms this is achieved, an experimental investigation has been conducted. This examines how modifications to the corner separation affect an $M=2.5$ oblique shock reflection. The nature of the flow field is studied using flow visualisation, pressure-sensitive paint and laser Doppler anemometry. The results show that the size and shape of central separation vary considerably when the onset and magnitude of corner separation changes. The primary mechanism explaining the coupling between these separated regions appears to be the generation of compression waves and expansion fans as a result of the displacement effect of the corner separation. This is shown to modify the three-dimensional shock structure and alter the adverse pressure gradient experienced by the tunnel floor boundary layer. It is suggested that a typical oblique SBLI in rectangular channels features several zones depending on the relative position of the corner waves and the main interaction domain. In particular, it has been shown that the position of the corner ‘shock’ crossing point, found by approximating the corner compression waves by a straight line, is a critical factor determining the main separation size and shape. Thus, corner effects can substantially modify the central separation. This can cause significant growth or contraction of the separation length measured along the symmetry line from the nominally two-dimensional baseline value, giving a fivefold increase from the smallest to the largest observed value. Moreover, the shape and flow topology of the centreline separation bubble is also considerably changed by varying corner effects.

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.

Shock control bumps offer the potential to reduce wave drag on transonic aircraft wings. However, most studies to date have only considered unswept flow conditions, leaving uncertain their applicability to realistic finite swept wings. This paper uses a swept infinite-wing model as an intermediate step, and it presents a computational study of the design drag performance of three-dimensional bumps. A new geometric parameter, termed bump orientation, is introduced and found to be crucial to the performance under swept flows. Classic shock control bumps aligned approximately with the local to freestream flow direction can offer drag reductions comparable to those from similar but unoriented devices in unswept flows, whereas badly misaligned bumps see severe performance degradation. For appropriately aligned classic bumps, the relationships between performance and selected geometric parameters (height, streamwise position, and isolation) are found to be somewhat similar to those observed in unswept studies.

© 2015 by Todd Davidson and Holger Babinsky. Published by the American Institute of Aeronautics and Astronautics, Inc.This paper reports recent work undertaken to investigate the interaction between normal shocks and laminar/transitional boundary layers. Laminar ow is desirable for its low skin friction, but problems are expected with shock interactions as the boundary layer will be prone to separation. Experiments have been performed on a at plate to explore the fundamental interaction without complex geometries and ow fields, and on a transonic aerofoil to research the in uence of the imposed pressure field. The boundary layer state before the shock on the at plate does not appear to significantly affect the downstream ow development, contrary to expectations, as the laminar separation is so small. On the aerofoil, a similar result is observed, with no change in the buffet onset Mach number. This implies that normal shock wave-laminar boundary layer interactions are not as detrimental as originally feared, and are in fact quite benign.

This paper presents results from a simplified experimental rig that aims to replicate the key physics of a transonic aircraft intake at high incidence for realistic altitudes. The equivalent flight conditions replicated are for free stream Mach numbers M∞ = 0.25 − 0.45 and incidence angles in the range α = 30 ± 5 degrees. Measurement techniques from the simplified two-dimensional geometric setup include; schlieren imaging, surface oilflows, pressure sensitive paint, pressure and temperature measurements. CFD models that look to predict this behaviour have a limited accuracy largely due to a lack of experimental validation data-this lends itself as one of the key motivations for the work presented in this paper. Currently there is good qualitative agreement between the experimental results and the initial computation(s).

A series of experiments have been conducted on a bleed hole array spanning the width of the Cambridge University Engineering Department supersonic wind tunnel at Mach numbers of 1.8 and 2.5. The wind tunnel was run with varying levels of suction, and the flow structure over the bleed array was subsequently mapped with a laser Doppler velocimetry system at a resolution of 0.25 hole diameters or better. The same wind-tunnel setup was simulated using the OVERFLOW Navier-Stokes equation solver. The information obtained was used primarily in qualitative comparisons of flow patterns. Overall good agreement was found in the definition of the expansion fan and barrier shock pattern produced by flow entering the normal holes, as well as three-dimensional flow patterns. Both studies agreed well in terms of measured mass flow rates, to within 1% of the boundary-layer mass flow. The presence of the barrier shock standing off the downstream edge of the bleed holes corresponded with a jet of upward flow, which may provide a mechanism for the generation of streamwise vortices.

The work here presents an experimental and computational study of a simplified intake bottom lip at incidence. The experiments are carried out in a small-scale blow-downwind tunnel.Computations are carried out in parallel to experiments using an in-house unstructured compressible flow solver. Evaluation of current Reynolds Averaged Navier-Stokes (RANS) methods are conducted and compared to experimental measurements. Qualitative results showgood agreement of experiments to the computations, although further analysis is required. Current computational strategies followa steady, quasi-two-dimensional approach, with fully structuredwall-resolved grids.

Results from wind tunnel experiments with three different geometry shock control bumps (SCBs) are presented and discussed. The primary aims of this study are to explore how 3-D SCBs generate stream-wise vorticity and to compare the flow features produced by different geometry 3-D SCBs. Results are presented in two parts: In part I, the flow development over a simple geometry baseline bump is described in detail with a focus on describing the evolution of vortical flow structures. In part II, results from tests with two different bumps (with subtle geometric variations from the baseline design) are presented and compared with those from the baseline bump. It is observed that the mechanism of vortex production for all 3-D SCBs tested is broadly in agreement with that reported in literature: Specifically, vortical flow structures are only observed downstream of the crest, on the SCB tail. Differences between the three SCB geometries tested are relatively subtle; increasing the width of the SCB tail reduces the extent of shock-induced separation at the crest, giving an anticipated reduction in viscous drag; steepening the SCB sides also reduces shock-induced separation at the crest (although to a lesser extent) and modifies the pressure profile on the off the bump such that the flow experiences lower span-wise pressure gradients. These findings suggest both geometric modifications considered may yield a benefit in performance for a 3-D SCB mounted on a transonic wing. © 2015, American Institute of Aeronautics and Astronautics Inc. All rights reserved.

This paper describes a numerical investigation into the optimal design of adaptive shock control bumps (SCB) for transonic wings. A multi-disciplinary approach to optimization is utilized, combining structural and aerodynamic analysis to ensure that optimal adaptive SCB do not exceed material constraints whilst maintaining aerodynamic qualities. It is found that adaptive SCB can perform to within 95% of the non-structurally constrained bumps but over a much wider ight envelope. Two dimensional single crest bumps have been shown to match the performance of table top bumps originating from three dimen- sional SCB. Total pressure recovery has been successfully used as a performance metric and will provide valuable comparisons to wind tunnel experiments with a prototype adaptive SCB.

An experimental study was performed to characterize the effect of flow control on the separated flow within a transonic diffuser. Mach 1.348 inflow at a Reynolds number of approximately 20 × 106 entered diffuser with a 6° slope to the floor. The baseline flow exhibited poor diffuser performance, reaching a static pressure rise of 46% of the predicted inviscid rise, and having a fully separated flow within the diffuser. The shock motion was not constrained, and Xrms was approximately equal to the baseline boundary layer thickness. Flow control was applied through 3 pairs of 3mm high ramped-vane vortex generators, or through suction across a distributed array of normal bleed holes with either 1mm or 2mm diameter. All forms of flow control were found to be effective at improving diffuser performance, with VGs increasing the static pressure to 63% of the inviscid rise, and suction through 1mm holes in the diffuser saw the static pressure increase to 80% of the inviscid pressure rise. Suction upstream of the diffuser entrance was found to produce a similar maximum pressure rise to that of VGs. Overall, VGs acted similar to boundary layer suction removing 15 - 20% of the floor boundary layer. Combinations of the two methods allowed for improvement similar to the high suction rates, while requiring 15% less of the boundary layer to be removed.

The effect of transition location on the interaction between normal shock waves and boundary layers is an area of study that has been enjoying a resurgence in interest recently, as operating regimes change and it becomes a phenomenon that may occur in flight. The problem is critical because a laminar boundary layer will typically separate at even a weak shock, with detrimental impact on the flow, when a turbulent boundary layer might remain attached. However, ahead of the shock the laminar boundary layer often preferable due to its lower drag. This investigation aims to explore the interaction between normal shocks and laminar/transitional flat plate boundary layers to gain insights into the fundamental flow physics of the problem. The flow is difficult to investigate experimentally, as although flat plate transition has been studied for many years, achieving the same phenomenon at high speeds and moderate Reynolds numbers is more challenging due to higher tunnel noise levels in comparison with flight or lower Reynolds number 'quiet tunnels'. This paper documents how an experiment has been developed with a laminar flat plate boundary layer and an incident normal shock wave. Transition has been triggered upstream of the shock, and some initial results of these transitional interactions are presented.

Two-dimensional and three-dimensional contour bumps are designed and optimized for substantial wave drag reduction for an un-swept natural laminar flow (NLF) wing (RAE5243 aerofoil section) at transonic speeds. An NLF aerofoil wing is chosen in this study, as shock control is more crucial for such wings due to the requirement of favourable pressure gradients on a Substantial part of the wing. For the validation purpose and to focus on the wave drag issues, the boundary layer is assumed to be fully turbulent from the leading edge. Key bump geometrical parameters including the maximum height, the length, and the crest position have been chosen for the parameterization of the two-dimensional and three-dimensional shock control bumps. For the three-dimensional bumps, an array of the contour bumps is installed spanwise on the transonic wing and their width and spanwise spacing are chosen as additional design parameters. Both the two-dimensional and the three-dimensional bump shapes are optimized using a discrete adjoint-based optimization method. The performance of the three-dimensional contour bumps are compared in detail with the similarly optimized two-dimensional bumps both at and around the design point. The results show that, for the NLF wing studied, the optimized three-dimensional bumps are as effective as the optimized two-dimensional bump in terms of total drag reduction at the given design point, despite the significant difference in their geometrical shapes. More importantly, in terms of the operational range for varying lift conditions for practical applications, the three-dimensional bumps outperform the two-dimensional bump by a substantial margin.

We analyse the currently popular vortex identification criteria that are based on point-wise analysis of the velocity gradient tensor. A new measure of spiralling compactness of material orbits in vortices is introduced and using this measure a new local vortex identification criterion and requirements for a vortex core are proposed. The inter-relationships between the different criteria are explored analytically and in a few flow examples, using both zero and non-zero thresholds for the identification parameter. These inter-relationships provide a new interpretation of the various criteria in terms of the local flow kinematics. A canonical turbulent flow example is studied, and it is observed that all the criteria, given the proposed usage of threshold, result in remarkably similar looking vortical structures. A unified interpretation based on local flow kinematics is offered for when similarity or differences can be expected in the vortical structures educed using the different criteria.

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.

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.

The evolution of a single hairpin vortex-like structure in the mean turbulent field of a low-Reynolds-number channel flow is studied by direct numerical simulation. The structure of the initial three-dimensional vortex is extracted from the two-point spatial correlation of the velocity field by linear stochastic estimation given a second-quadrant ejection event vector. Initial vortices having vorticity that is weak relative to the mean vorticity evolve gradually into omega-shaped vortices that persist for long times and decay slowly. As reported in Zhou, Adrian & Balachandar (1996), initial vortices that exceed a threshold strength relative to the mean flow generate new hairpin vortices upstream of the primary vortex. The detailed mechanisms for this upstream process are determined, and they are generally similar to the mechanisms proposed by Smith et al. (1991), with some notable differences in the details. It has also been found that new hairpins generate downstream of the primary hairpin, thereby forming, together with the upstream hairpins, a coherent packet of hairpins that propagate coherently. This is consistent with the experimental observations of Meinhart & Adrian (1995). The possibility of autogeneration above a critical threshold implies that hairpin vortices in fully turbulent fields may occur singly, but they more often occur in packets. The hairpins also generate quasi-streamwise vortices to the side of the primary hairpin legs. This mechanism bears many similarities to the mechanisms found by Brooke & Hanratty (1993) and Bernard, Thomas & Handler (1993). It provides a means by which new quasi-streamwise vortices, and, subsequently, new hairpin vortices can populate the near-wall layer.

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.

An experimental study of several vortex generators with heights well less than that of the boundary-layer (“micro-” VGs) has been conducted in a supersonic wind tunnel to better understand vortex strength and development downstream of the device. VGs known as micro-ramps and common-flow up vane pairs of two heights were tested at Mach 1.5. The static pressure distribution and surface flow topology on the tunnel floor near the devices have been imaged, and two components of velocity in the wake have been measured using LDA. These data in combination with schlieren photographs give new insight into the flow structure generated by these devices. The development of the primary vortices generated by these devices is examined and the vortices’ evolution, strength, and trajectory are estimated as they travel downstream of the device. These devices are seen to behave similarly to those tested previously by others in subsonic flows.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.