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# Comparison of experimental and computational shock structures for three bumps at o↵-design condition: (a) HSCB; (b) wedge bump; (c) exSCB

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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 b...

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A comprehensive experimental investigation of helicopter blunt fuselage drag reduction using active flow control is being carried out within the European Clean Sky program. The objective is to demonstrate the capability of several active flow technologies to decrease fuselage drag by alleviating the flow separation occurring in the rear area of som...

## Citations

... The pressure in the separated flowfield varies from 29 − 57 kPa; this range of pressures is sufficiently large to provide reliable measurements (Sajben 1993). In order to determine the values of A(T) and B(T) in the Stern-Volmer relation, in-situ calibration is performed using five 0.3 mm diameter static pressure taps connected to a differential pressure transducer (error: ±1% ) (Colliss et al 2016). This calibration enables absolute pressure values on the target surface to be extracted from the measured light intensity. ...

... The streamwise and floor-normal flow velocities, u and v, respectively, are measured by two-component laser Doppler velocimetry (LDV). The flow is seeded with paraffin in the settling chamber; previous measurements of particle lag through a normal shock have placed the seeding droplet diameter in the range 200 − 500 nm (Colliss et al 2016). The measured velocities have an error of 1% and 14% for u and v, respectively; there are contributions from the number density of seeding particles and from the laser optics. ...

Many supersonic wind tunnel experiments investigate shock–boundary-layer interactions by measuring the response of tunnel wall boundary layers to an incident shock wave. To generate the supersonic flow, these facilities typically use two-dimensional contoured converging–diverging nozzles which can be arranged in two different ways. One configuration is symmetric about the centre height, whereas this symmetry plane defines the tunnel floor in the other asymmetric arrangement. In order to determine whether these nozzle configurations, which are widely thought to be equivalent, can influence experiments on shock–boundary-layer interactions, two different nozzle geometries are compared with one another in a single facility with rectangular cross section. For each setup, a full-span 8-degree wedge introduces an oblique shock to a Mach 2.5 flow. The two setups exhibit quite dissimilar behaviour, both in the corner regions and on the tunnel’s centre span, with a difference in central separation length of as much as 35% suggesting that nozzle geometry can have a profound impact on these interactions. The observed behaviour is caused by known secondary flows in the sidewall boundary layers which are driven by vertical pressure gradients in the nozzle region. The subsequent impact on the response of the floor boundary layer is consistent with expectations based on local flow momentum affecting corner separation size and on the displacement effect of this corner separation influencing the wider flow.
Graphical abstract

... With this device the quasi-normal shock wave is replaced by a λ-shaped shock wave, with the working principle extensively described by Bruce and Colliss (2015). Colliss et al. (2016); Ogawa et al. (2008) have confirmed that an array of 3D SCBs is more efficient than a 2D bump configuration (which spans along the full span of the model as described in detail by Zhang et al. (2021)) in view of the streamwise vortices developing from the tail of the bumps. Applications of the use of SCBs for controlling buffet are discussed in the numerical studies of Mayer et al. (2018) and Geoghegan et al. (2020), which confirmed the dependence of the performance of a SCB on its size, shape and position relative to the shock. ...

This experimental study investigates the possibility of controlling transonic buffet by means of a trailing edge flap with an upward deflection (referred to as “upper trailing edge flap”, or: UTEF). Different geometries (straight and serrated) and dimensions of UTEFs (with heights ranging between 1 and 2% of the chord) have been studied with respect to their impact on the buffet behavior. The effectiveness of the UTEFs has been investigated with schlieren and particle image velocimetry (PIV) in the transonic-supersonic wind tunnel of TU Delft at Ma = 0.70, α = 3.5°. The schlieren results demonstrated the efficacy of the use of UTEFs for reducing the range of the buffet oscillations when the height of the UTEF was equal to at least 1.5%c. This result was corroborated by a flow characterization with PIV data and which highlighted that, in presence of a control system, not only the shock oscillation range is reduced but also the intensity of the separated area pulsation. The use of serrated UTEFs, despite having an effect on the local flow field, was found to be ineffective in alleviating buffet oscillations. The adoption of the best behaving UTEF configuration (straight 2%c UTEF) proved to only slightly alter the circulation value compared to the clean configuration, while it also proved to be effective in an off-buffet condition (Ma = 0.74 and α = 2.5°).

... In the US, the N+3 goal proposed by NASA is to reduce Nitrogen oxides (NOx) emission by up to 80% in the landing-take-off process and reduce fuel burn by 60% for an airliner entering service in 2030-2035 [3]. To achieve these objectives, a number of technologies, such as shock control [4][5][6][7], laminar flow control [8][9][10][11][12][13][14], turbulent drag reduction [15][16][17][18], as well as novel aircraft concepts, such as BWB or hybrid wing body (HWB) [19], 'double-bubble' [20], truss-braced wing (TBW) [21] and box-wing [22], have been proposed and investigated to explore a better aerodynamic performance. However, there are significant challenges in applying these technologies mentioned above on future aircraft, especially in terms of practical application. ...

Effective control of aerodynamic loads, such as maneuvering load and gust load, allows for reduced structural weight and therefore greater aerodynamic efficiency. After a basic introduction in the types of gusts and the current gust load control strategies for aircraft, we outline the conventional gust load alleviation techniques using trailing-edge flaps and spoilers. As these devices also function as high-lift devices or inflight speed brakes, they are often too heavy for high-frequency activations such as control surfaces. Non-conventional active control devices via fluidic actuators have attracted some attention recently from researchers to explore more effective gust load alleviation techniques against traditional flaps for future aircraft design. Research progress of flow control using fluidic actuators, including surface jet blowing and circulation control (CC) for gust load alleviation, is reviewed in detail here. Their load control capabilities in terms of lift force modulations are outlined and compared. Also reviewed are the flow control performances of these fluidic actuators under gust conditions. Experiments and numerical efforts indicated that both CC and surface jet blowing demonstrate fast response characteristics, capable for timely adaptive gust load controls.

... Limitations to the optical access in this facility prevent a three-component system from being installed and so only the streamwise and vertical components of velocity, and , can be obtained. Previous measurements in this facility of particle lag through a normal shock have placed the seeding droplet diameter in the range 200 -500 nm [11]. The error in measured velocities due to the finite number density of seeding particles and due to the laser optics is 2% for and 19% for . ...

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.

... Since then, several types of shock control bump, including contour and wedge-shaped bumps, have been extensively investigated by both computational and experimental studies [2,3,[8][9][10][11][12][13][14][15]. The dominant drag reduction mechanism is either the weakening of the shock wave or the turning of the shock into continuous compression waves, resulting in the reduction of the wave drag. ...

... Wong et al. [12] and Ogawa et al. [13] observed the existence of streamwise vortices after 3D ramp bumps. Later, the vortical structures were further investigated by Colliss et al. [15] in a joint experimental and computational study. ...

... The vortical flows appear as the by-products of their sharp edges. The vortices are usually weak [15,27,29] and hard to control without affecting the effectiveness of shock control as shown in the study by Mayer et al. [29]. As this type of 3D bumps is designed for shock control, their capability in suppressing flow separation is unknown. ...

The 3D contour bump has been integrated with vane-type vortex generators to weaken the shock wave and suppress potential flow separation at higher transonic speeds. The integrated vortex-generating shock control bump is parameterized as a whole, in a manner that the contour bump is designed for shock control and the vortex-generating fins are integrated to the contour bump for separation control. This allows the balance between the two factors affecting the total wing drag. The optimization studies based on the Reynolds-averaged Navier–Stokes equations have been carried out to find the optimal vortex-generating bump designs at the given design conditions. Single-point and multipoint global optimizations are carried out to search the optimum parameters at the corresponding design points. It is found that the vortex-generating bump can further reduce the total drag at the higher transonic Mach numbers due to the alleviation of after-bump streamwise separation by a pair of counter-rotating vortices generated. The inclusion of the vortex generators in the bump design allows for a better balance of the design in controlling both the shock strength and the flow separation after the bump at higher Mach numbers. A multipoint optimization leads to a robust vortex-generating bump design for a range of Mach numbers. The distinctive vortical flow structures induced by the vortex-generating bump are highlighted and discussed.

... The onset of corner separation and its magnitude is varied by controlling the corner effects by suction and creating an obstacle in the flow. For smaller corner separation, they found surface topology corresponding to the owl face of the first kind, as reported first by Perry and Hornung [29]; and for more significant corner separation, they found owl face of the second kind, as reported by Colliss et al. [30]. Nevertheless, all the cases had three-dimensionality. Wang et al. [31] made an essential contribution in explaining the three-dimensional nature of the separation bubble. ...

The influence of a finite-width shock generator on the incident shock in the impinging shock-wave/boundary-layer interaction is investigated using different widths of the shock generator. The shock-wave generator widths are kept at 80, 50, and 20% of the total width of the bottom wall. A highly three-dimensional incident oblique shock wave is observed. Experimental and numerical studies are used to understand the nature of three-dimensional shock downstream of the leading edge of the shock generator. The incident shock curved downstream in the streamwise and spanwise directions, and the curvature increased with the decreasing width of the shock generator. The local strength of the incident shock at the bottom wall varied significantly due to this curvature for each of the cases. Consequently, even for shock generators with the same flow deflection, the shock-wave/boundary-layer interaction nature transitioned from separated to unseparated. An explanation and a model for the three-dimensionality of the incident shock are presented, and numerical simulations validate the model. The model shows that the shape of the incident shock is a function of the freestream Mach number, the flow deflection angle, and the width of the shock generator.

... Approximately 10% drag reduction was achieved in the drag-rise region in their experiment. Some fundamental understanding of the flow physics for shock control bumps can be found in Barbinsky et al. [27] for different ramp bumps mounted on the floor of a supersonic wind tunnel. In order to extend the laminar flow region and meanwhile alleviate the shock wave, Tang et al. ...

At a transonic condition, the design of a natural laminar flow (NLF) wing is challenging because the extension of the laminar flow needs to be finely balanced with the potential wave drag increase. To achieve this balance, it is proposed to unlock the wing sweep and introduce a three-dimensional (3-D) contour shock-control bump (SCB) in the optimization of the NLF infinite swept wing aiming at total drag reduction. A 3-D Reynolds-averaged Navier–Stokes flow solver is extended to incorporate laminar–turbulent transition prediction due to streamwise and crossflow instabilities. The flow solver is integrated in a gradient-based optimization framework. The transition criteria including streamwise and crossflow instabilities are coupled in the sensitivity calculation using a discrete-adjoint solver. Transonic design optimization of the sectional profile and wing sweep angle is first conducted at a Mach number of 0.78. The optimization managed to alleviate the shock wave, reducing the pressure drag while it failed to extend the laminar flow. Furthermore, a combined optimization of the wing with a parameterized SCB at the same condition is carried out. The optimized wing with the SCB features a long favorable pressure gradient region and a low-sweep angle, which allows for a large proportion of laminar flow without significant pressure drag. The shock wave is controlled by the 3-D bump, and the SCB has little effect on the upstream flow.

... The flow is seeded with paraffin in the settling chamber. Previous measurements in this facility of particle lag through a normal shock have placed the seeding droplet diameter in the range 200 − 500 nm [11]. The measured velocities have an error of 2%, due to the finite number density of seeding particles and due to the laser optics. ...

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.

... They showed that 3D bump configurations can be more robust than 2D bump designs through adjoint-based design optimisation. Collins et al. (22) tested shock control bumps in a wind tunnel, and analysed the performance of shock control bumps at off-design conditions. They noted that, at off-design conditions, even smooth bumps can create vortical flow between the individual bumps and that large, separated regions occur. ...

Shock control bumps can be used to control and weaken the shock waves that form on engine intakes at high angles of attack. In this paper, it is demonstrated how shock control bumps applied to an engine intake can reduce or eliminate shock-induced separation at high incidence , and also increase the incidence at which critical separation occurs. Three-dimensional Reynolds-average Navier-Stokes (RANS) simulations are used to model the flow through a large civil aircraft engine intake at high incidence. The variation in shock strength and separation with incidence is first studied, along with the flow distribution around the nacelle. An optimisation process is then employed to design shock control bumps that reduce shock strength and separation at a fixed high incidence condition. The bump geometry is allowed to vary in shape, size, streamwise position and circumferential direction around the nacelle. This is shown to be key to the success of the shock control geometry. A further step is then taken, using the optimisation methodology to design bumps that can increase the incidence at which critical separation occurs. It is shown that, by using this approach, the operating range of the engine intake can be increased by at least three degrees.

... The flow is seeded with para n in the settling chamber. Previous measurements in this facility of particle lag through a normal shock have placed the seeding droplet diameter in the range 200 500 nm [7]. The measured velocities have an error of 2%, due to the finite number density of seeding particles and due to the laser optics. ...

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.