In this paper, a non-intrusive pressure measurement scheme based on particle image velocimetry (PIV) is presented for the complex supersonic flows with intense shock systems, by elaborately combining the MacCormack method, the streamline-based method, and the spatial integration in conservative form. According to the detailed analyses of flow structures, the pressure fields are well reconstructed by the proposed scheme for the two typical shock-wave/boundary-layer interactions containing regular and Mach reflections, which are induced by the relatively strong oblique shock waves generated by the wedges of 21° and 17° in the freestreams of Mach 2.5 and 2.0, respectively. Based on the theoretical solutions by oblique shock relationship, free interaction theory, and shock polar analysis, this pressure reconstruction scheme is completely validated to effectively suppress the propagation of PIV velocity error to the pressure field and the accumulation of reconstructed pressure error behind the strong shock wave. Compared with the literature presently, this work would be the most challenging application of PIV-based pressure measurement to such complex supersonic flows with intense shock reflections, large oscillations, wide speed ranges, and various compressible flow structures. These good results could confirm the feasibility and high accuracy of the proposed reconstruction scheme and may greatly promote its applications in academic research and engineering test for supersonic flows in the future.

@en

The discovery of vast subsurface oceans hidden under kilometers of ices on icy moons in our Solar System has sparked worldwide interests in ascertaining their potential habitability. In the case of Saturn’s moon Enceladus, supersonic plumes of water vapour and icy grains have been observed by the Cassini mission spewing from the surface, giving us indirect knowledge of the composition of this subsurface ocean. The exact mechanisms of the plumes however, and their effect on the composition of the ejected matter has yet to be clearly understood. The focus of this study is to experimentally investigate physical characteristics of the plumes located at the South Polar Terrain (SPT) of Enceladus. Using facilities at TU Delft faculty, we simulate experimentally the topology of the ice crevasses and the subsurface ocean with a narrow channel mounted atop a liquid water reservoir placed inside a vacuum chamber. We inquire upon the dependence of the channel walls temperature on the plume’s exhaust velocity. Using a straight channel, our results show that colder wall temperatures enable a saturated water vapour flow with a minima 1.5-3 % solid fraction and vent velocities reaching around 400-500 m/s. The data ranges for velocities and solid fraction extrapolated from the Cassini data (550-2000 m/s and 7-70 %) thus cannot be explained by straight channel models. Using a channel with an expansion ratio of 1.73 however, the measured supersonic plume velocity becomes comparable to some of the in situ Mach number determined at Enceladus. Using a method based on the energy conservation law, it is possible to extrapolate from our experimental data some plausible geometries of the ice crevasses of Enceladus. This work lays the ground work for a coming comprehensive parametric study of the channel geometry and its effect on exhaust Mach number, temperature and solid fraction.@en

At cruise altitudes, the Reynolds number may become sufficiently low to allow a laminar boundary layer to persist on the suction side of a transonic fan blade up to the shock-wave/boundary-layer interaction (SBLI). In such a transitional SBLI with sufficiently large shock-induced separation, a shock oscillation mechanism occurs, with the source at the upstream growth (until a critical length) and natural suppression (through shear layer instabilities) of the separation bubble. The oscillation cycle is characterized by a temporarily vanishing upstream laminar part of the separation bubble. The suppression of this laminar part is accompanied by downstream advection of turbulence and subsequent entrainment into the bulk separation bubble, affecting the reflected shock movement. In order to study the spatial behavior of the mechanism, particle image velocimetry of a highly separated transitional oblique SBLI at Mach 2.3 is conducted in the high speed aerodynamics laboratory of Delft University of Technology. Statistical quantities, including root mean square velocity fluctuations, phase averages, and spatial modes from proper orthogonal decomposition, are investigated. The entrainment strength varies depending on the phase of the oscillation, and the turbulent shear layer is not fully developed. The main growth and shrinking mode of the separation bubble were extracted, which affects the slip line region size and shock position. Secondary modes that affect the shear layer undulation and upstream effects were also extracted. The study provides quantitative analyses of an important shock oscillation type, with the focus on capturing the separation bubble size variation and upstream effects.

@en

Modern launcher configurations, often characterized by a larger diameter in the payload fairing than the rest of the launch vehicle (hammerhead configuration), face significant challenges during transonic operations, given their susceptibility to flow separation and intense pressure fluctuations. This experimental study aims to reconstruct pressure from PIV under transonic conditions using Taylor's hypothesis. The focus is on the ESA VEGA-E hammerhead launcher model, investigated in the transonic regime at Ma=0.8 and at α=0°. Initially, the methodology of the pressure reconstruction algorithm is described and the validity of Taylor's hypothesis is established through validation with numerical data. Subsequently, the experimental data are first used to provide a general characterization of the flow phenomenon, highlighting key features such as an oscillating shockwave, separated and reattached flows. Later, the reconstruction of pressure from velocity data is compared to unsteady pressure transducer data, for instantaneous, average and standard deviation values, obtaining a very good agreement (ΔCpAvg~0.01-0.02 and ΔCpStd~0.02). The experimental set-up is also designed to showcase the impact of neglecting the out-of-plane velocity component on pressure reconstruction, showing comparable data for CpAvg, while larger discrepancies in terms of CpStd in particular in the separated area region.@en

Supersonic plumes of water vapour and icy particles have been observed by the Cassini spacecraft during several flybys over Enceladus. These plumes originate from the Tiger Stripes located in the South Polar Terrain (SPT), and indicate the presence of a subsurface ocean under the icy crust which is salty and contains complex organic molecules. Other characteristics of the plumes, such as the vent temperature, mass flow rate, velocity and mass fraction of icy particles can be used to determine the conditions in the channel, linking the subsurface ocean to the icy surface. In this paper, we developed a fluid dynamics model that accounts for nucleation, particle growth, wall accretion and sublimation. The channel behaves similarly to a converging–diverging nozzle, which forms supersonic plumes due to a pressure difference between the reservoir where the subsurface ocean is located and the exosphere. The geometry of the channel and its evolution with accretion of gas and sublimation of ice are studied to reproduce the characteristics of the plumes observed by Cassini. We first performed a parameter study on the channel geometry to determine how it influences the plumes’ velocity, solid fraction and exit temperature. Our results show that the size of the icy particles is primarily dependent on the length of the channel, indicating that large particles (∼75μm) must originate from within a kilometer below the surface, while smaller particles (∼3μm) can originate from only hundreds of meters below the surface. We further show that the velocity of the flow, exit temperature and nucleation depend directly on the exit-to-throat size ratio. We find that the channel geometry evolves within a few tens of hours until an equilibrium is reached, when considering the accretion of gas to the walls, or sublimation of ice from the walls. As the channel closes due to accretion, the flow becomes thinner, which in turn reduces accretion. After around 70 h, the accretion is sufficiently slowed such that the geometry does not evolve anymore. This equilibrium geometry produces higher Mach numbers and a larger particle size and solid fraction compared to the initial geometry.

@en

Wall-resolved large-eddy simulations (LES) are performed to investigate Reynolds number effects in supersonic turbulent boundary layers (TBLs) at Mach 2.0. The resulting database covers more than a decade of friction Reynolds number Re_τ, from 242 to 5554, which considerably extends the parameter range of current high-fidelity numerical studies. Reynolds number trends are identified on a variety of statistics for skin-friction, velocity and thermodynamic variables. The efficacy of recent scaling laws as well as compressibility effects are also assessed. In particular, we observe the breakdown of Morkovin's hypothesis for third-order velocity statistics, in agreement with previous observations for variable-property flows at low Mach number. Special attention is also placed on the size and topology of the turbulent structures populating the TBL, with an emphasis on the outer-layer motions at high Reynolds number. The corresponding streamwise spectra of streamwise velocity fluctuations show a clear separation between inner and outer scales, where energetic peaks are found at streamwise wavelengths of λ_x⁺≈700 and λ_x/δ₀≈6. The spanwise spacing of the outer-layer structures, in turn, is found to be insensitive to the Reynolds number and equal to ∼0.7δ₀. It is also found that the integral length-scales in spanwise direction for the temperature, streamwise and spanwise velocity fields appear to progressively collapse with increasing Reynolds number. The modulating influence that the outer-layer structures exert on the near-wall turbulence is also clearly visible in many of the metrics discussed. In addition, the present LES data is further exploited to assess the Re_τ-sensitivity of uniform momentum regions in the flow. We find that the resulting probability density function of the number of zones as well as its evolution with Re_τ agrees well with incompressible data. This suggests that uniform zones, which have been associated with outer-layer dynamics, are not strongly influenced by compressibility at the considered Mach number.

@en

In this experimental study, the effect of three-dimensional shock control bumps (SCB) on impinging-reflecting shock wave/boundary layer interactions (SWBLI) is investigated as a passive control method. The aim is to develop an understanding of the influence of such devices on the interaction structure by studying the position of the bump with respect to the shock-impingement location. The experiments were conducted in the ST-15 wind-tunnel at the Delft University of Technology for fully developed turbulent boundary layer conditions with Mach number of 2. The effectiveness was assessed by examining the size of the separated flow region, as well as the downstream boundary layer velocity profile. Stereo-PIV was employed as the main diagnostic method to characterise the three-dimensional flow field. Additionally, the effect of shock impingement location on the unsteady interaction dynamics was examined by analyzing the separation size and reflected shock foot position in the instantaneous flow. The investigation revealed the significance of the shock impingement location in terms of control effectiveness. Nevertheless, placement of the bump in the interaction region substantially decreased the probability of flow separation even in off-design conditions when compared to the uncontrolled interaction.@en

We investigate Reynolds number effects in strong shock-wave/turbulent boundary-layer interactions (STBLI) by leveraging a new database of wall-resolved and long-integrated large-eddy simulations. The database encompasses STBLI with massive boundary-layer separation at Mach 2.0, impinging-shock angle 40^◦ and friction Reynolds numbers Re_τ 355, 1226 and 5118. Our analysis shows that the shape of the reverse-flow bubble is notably different at low and high Reynolds number, while the mean-flow separation length, separation-shock angle and incipient plateau pressure are rather insensitive to Reynolds number variations. Velocity statistics reveal a shift in the peak location of the streamwise Reynolds stress from the separation-shock foot to the core of the detached shear layer at high Reynolds number, which we attribute to increased pressure transport in the separation-shock excursion domain. Additionally, in the high Reynolds case, the separation shock originates deep within the turbulent boundary, resulting in intensified wall-pressure fluctuations and spanwise variations associated with the passage of coherent velocity structures. Temporal spectra of various signals show energetic low-frequency content in all cases, along with a distinct peak in the bubble-volume spectra at a separation-length-based Strouhal number St_Lsep ≈ 0.1. The separation shock is also found to lag behind bubble-volume variations, consistent with the acoustic propagation time from reattachment to separation and a downstream mechanism driving the shock motion. Finally, dynamic mode decomposition of three-dimensional fields suggests a Reynolds-independent statistical link among separation-shock excursions, velocity streaks and large-scale vortices at low frequencies.

@en

Recent numerical studies have suggested the potential of substrates with streamwise-preferential permeability to reduce drag in turbulent boundary layers. Such a substrate is theorized to facilitate relaxation of the no-slip condition and thereby reduce the skin friction. So far, these beneficial effects have not been demonstrated experimentally yet and therefore the scope of this work is to present this concept in air flow where the substrate geometry satisfies the theoretical permeability requirements for an expected reduction in drag. For this, a three-dimensional-printed structure with anisotropic permeability (φxz=2.7, φxy=3.9) and small pores (s≈250μm), akin to an acoustic liner, was developed. The substrate was investigated using direct force measurements and 2D-2C PIV in the range of U∞≈5-35 ms-1, corresponding to frictional Reynolds numbers of Reτ≈430-1960. Results show an increase in drag of 0%<ΔCD<8% and, while contrasting the model predictions, this agrees with DNS data on structures with similar geometric properties when using the inverse wall-normal Forchheimer coefficient, or inertial permeability, as the equivalent roughness parameter. Hence the present results constitute the first experimental evidence that this is the governing property for the drag behavior of acoustic liners. The absence of the predicted beneficial flow modulation effects is attributed to the investigated substrate not strictly satisfying the theoretical framework assumptions on characteristic length scales. However, to expand beyond this structural limitation, we analytically derive that, for realistic, geometrically resolved cases, this length scale mismatch is unavoidable and thereby render it unfeasible to model the substrate as a continuum for the virtual-origin approach. We expect that translating the abstraction of substrates with streamwise-preferential permeability into physical realisations relevant for practical applications would result in structures very similar to riblets.

@en

The dynamic coupling between a Mach 2.0 shock-wave/turbulent boundary-layer interaction (STBLI) and a flexible panel is investigated. Wall-resolved large-eddy simulations are performed for a baseline interaction over a flat-rigid wall, a coupled interaction with a flexible panel, and a third interaction over a rigid surface that is shaped according to the mean panel deflection of the coupled case. Results show that the flexible panel exhibits self-sustained oscillatory behavior over a broad frequency range, confirming the strong and complex fluid-structure interaction (FSI). The first three bending modes of the panel oscillation are found to contribute most to the unsteady panel response, at frequencies in close agreement with natural frequencies of the mean deformed panel rather than those for the unloaded flat panel. This highlights the importance of the mean panel deformation and the corresponding stiffening in the FSI dynamics. The time-averaged flow shows an enlarged reverse-flow region in the presence of mean surface deformations. The separation-shock unsteadiness is enhanced due to the panel motion, leading to higher wall-pressure fluctuations in the coupled interaction. Spectral analysis of the separation-shock location and bubble-volume signals shows that the STBLI flow strongly couples with the first bending mode of the panel oscillation. This is further confirmed by dynamic mode decomposition of the flow and displacement data, which reveals variations in the reverse-flow region that follow the panel bending motion and appear to drive the separation-shock unsteadiness. Low-frequency modes that are not associated with the fluid-structure coupling, in turn, are qualitatively similar to those obtained for the rigid-wall interactions, indicating that the characteristic low-frequency unsteadiness of STBLI coexists with the dynamics emerging from the fluid-structure coupling. Based on the present results, unsteady FSIs involving STBLIs and flexible panels are likely to accentuate rather than mitigate the undesirable features of STBLIs.

@en

For the largest wind turbines currently designed, when operating at rated power and at high wind speeds, the tip airfoils can experience large negative angles of attack. For these conditions and in combination with turbulence, the airfoils are at risk of reaching locally supersonic flow, even at low free-stream Mach numbers. The possibility of shock wave formation and its consequences endangers the lifetime of these largest rotating machines ever built. So far only numerical analyses of this challenge have been attempted with significant modelling uncertainty. Here, for the first time, a wind turbine airfoil (the FFA-W3-211, used at the blade tip of the IEA 15MW reference wind turbine) is studied under transonic conditions using experimental techniques. Schlieren visualization and Particle Image Velocimetry were employed for free-stream Mach numbers of 0.5 and 0.6 and various angles of attack. It was shown that calculations based on isentropic flow theory and compressibility corrections were able to predict the situations where supersonic flow occurred. However, they could not predict the frequency of occurrence and whether shock waves were formed. In conclusion, an unsteady characterization of such airfoil behavior in transonic flow seems to be warranted.

@en

Hammerhead launcher configurations, characterized by a larger diameter in the payload fairing than the rest of the launch vehicle, face substantial challenges during transonic operations due to their susceptibility to flow separation and intense pressure fluctuations. This experimental study investigates the influence of the nose and boat-tail geometry on the flow around hammerhead configurations in the transonic regime (Ma=0.7-0.8) and for various angles of attack (α=0-4°). To gain a general understanding of the main flow features, such as shockwave formation, separated flow in the boat tail region, and flow reattachment, oil flow and schlieren visualizations were employed. Schlieren visualizations were also utilized to characterize the level of unsteadiness in these regions. Additionally, particle image velocimetry was employed to quantify variations in the velocity field. The study's findings reveal an optimization of flow performance in the presence of a bi-conic nose, attributed to the creation of two-shockwave structures with relatively low intensity. This is in contrast to the ogive and conic noses, which exhibit a single, more detrimental shockwave structure (with the conic nose being the least favorable configuration). The investigation into different boat tail angles indicates that adopting low-angle boat tails (5° and 15° compared to 34°) leads to a noticeable reduction in the separated area, albeit associated with an increase in the range of oscillation of the shockwave structures.

@en

We present an experimental realisation of spatial spanwise forcing in a turbulent boundary layer flow, aimed at reducing the frictional drag. The forcing is achieved by a series of spanwise running belts, running in alternating spanwise direction, thereby generating a steady spatial square-wave forcing. Stereoscopic particle image velocimetry in the streamwise–wall-normal plane is used to investigate the impact of actuation on the flow in terms of turbulence statistics, drag performance characteristics, and spanwise velocity profiles, for a non-dimensional wavelength of λ_x⁺=397. In line with reported numerical studies, we confirm that a significant flow control effect can be realised with this type of forcing. The scalar fields of the higher-order turbulence statistics show a strong attenuation of stresses and production of turbulence kinetic energy over the first belt already, followed by a more gradual decrease to a steady-state energy response over the second belt. The streamwise velocity in the near-wall region is reduced, indicative of a drag-reduced flow state. The profiles of the higher-order turbulence statistics are attenuated up to a wall-normal height of y⁺≈100, with a maximum streamwise stress reduction of 45% and a reduction of integral turbulence kinetic energy production of 39%, for a non-dimensional actuation amplitude of A⁺=12.7. An extension of the classical laminar Stokes layer theory is introduced, based on the linear superposition of Fourier modes, to describe the non-sinusoidal boundary condition that corresponds to the current case. The experimentally obtained spanwise velocity profiles show good agreement with this extended theoretical model. The drag reduction was estimated from a linear fit in the viscous sublayer in the range 2≤y⁺≤5. The results are found to be in good qualitative agreement with the numerical implementations of Viotti et al. (Phys Fluids 21, 2009), matching the drag reduction trend with A⁺, and reaching a maximum of 20%. Graphical abstract: (Figure presented.)

@en

Chevron-shaped protrusions have been proposed in the literature for turbulent skin friction reduction. However, there is no consensus on the performance of this passive flow control technique; both an increase and a decrease in drag have been observed in previous studies. There is also no experimental evidence to support the working mechanism behind the drag reduction effect that has been postulated in the literature. In this study, direct force measurements were used to replicate experiments from the literature and, in addition, were used to test new array configurations to characterise the effect of individual design parameters on drag performance. A total of 23 different protrusion configurations were investigated in a turbulent boundary layer flow. In addition to the integral force measurements, particle image velocimetry was used to measure wall-parallel velocity fields in order to extract the statistical sizing and energy of the near-wall cycle turbulence. All configurations increased the drag between 2% and 10% for a friction Reynolds number of 1700. The drag reduction reported in the literature could not be replicated; however, these findings agreed with an experimental and numerical study that reported drag increase. The trend observed in the low-speed streak spacing from the PIV experiments was consistent with that observed in the balance data. Nevertheless, no evidence was found to support the working mechanism proposed in the literature. These results cast doubt on the proposed drag reduction potential of chevron-shaped protrusions. In the authors’ view, the results of this study strengthen previous conclusions regarding their minor increase in drag. Future studies to further approach a consensus are proposed.

@en

Boundary-layer instability on a rotating cone induces coherent spiral vortices that are linked to the onset of laminar–turbulent transition. This type of transition is relevant to several aerospace systems with rotating components, e.g., aeroengine nose cones. Because a variety of options exist for the nose-cone shapes, it is important to know how their shape affects the boundary-layer transition phenomena. This study investigates the effect of varying cone angle on the boundary-layer instability on rotating cones facing axial inflow. It is found that increasing cone angle has a stabilizing effect on the boundary layer over rotating cones in axial inflow. The parameter space of Reynolds number Re _l and local rotational speed ratio S is experimentally explored to find the spiral vortex growth on rotating cones of half angle ψ 22.5°, 45°, and 50°. The previously addressed cases of ψ 15° and 30° are also revisited. Increasing half-cone angle is found to have a stabilizing effect on the boundary layer on the rotating cones with ψ ≲ 45°; i.e., the spiral vortex growth is delayed to higher Re _l and S. This effect diminishes when the half-cone angle increases from ψ 45° to 50°. The spiral vortex angle ϵ decreases with increasing rotational speed ratio S for all the investigated cones, irrespective of the half-cone angle. However, the instability on the broader cones is found to induce shorter azimuthal wavelengths.

@en

The interaction between a shock wave and a boundary layer is a topic of primary relevance in high-speed aerodynamics, as it may deteriorate the vehicle performance, and can even lead to structural damage. Over the years, many researchers have investigated various control techniques to mitigate the detrimental effects of such shock wave boundary layer interactions (SWBLI). In this experimental study the effect of shock control bumps (SCB) on oblique shock wave/boundary layer interactions is investigated, notably the influence of the ramp section of the bump. To study the effect, varying bump geometries were designed with 5 different ramp angles while the maximum crest height is kept constant. The experiments were conducted in the ST-15 wind-tunnel at the Delft University of Technology for fully developed turbulent boundary layer conditions with Reθ of 21.8 103 and freestream Mach number of 2.0. The effectiveness is assessed from the size of the separated flow region, as well as the downstream boundary layer velocity profile. For this, PIV is employed as the main diagnostic method to characterise the flow field. In addition to this, high-speed Schlieren and oil flow measurements were performed to asses the effect of the SCB on the overall interaction structure. @en

Wall-resolved large-eddy simulations (LES) are carried out to investigate the aeroelastic coupling between a Mach 2.0 impinging shock-wave/turbulent boundary-layer interaction (STBLI) and a flexible thin-panel. After the initial transient, the panel exhibits self-sustained oscillatory behavior with varying oscillation amplitude, confirming the strong and complex dynamic coupling over a broad frequency range. The first three bending modes of the panel oscillation are found to contribute most to the unsteady panel response. The observed modal frequencies are in close agreement with natural frequencies of the pre-stressed panel, which differ significantly from the natural frequencies of the unloaded flat panel. This highlights the importance of the mean panel deformation and the corresponding stiffening in the fluid-structure interaction (FSI) dynamics. Mean-flow shows an enlarged reverse-flow region compared to a flat rigid-wall STBLI at the same flow conditions. The separation shock is also located further upstream in the coupled case, and wall-pressure fluctuations have a higher peak at the separation-shock foot. Spectral analysis of wall-pressure, separation-shock location and bubble-volume signals indicates that the STBLI flow resonates with the panel oscillation, primarily at the first bending frequency. This is further confirmed by sparsity-promoting dynamic mode decomposition of the flow and displacement data, which identifies this frequency as the most dominant and successfully isolates the associated FSI dynamics. Based on present results, it is clear that dynamic FSI involving STBLI and flexible panels accentuates the undesirable features of STBLI.@en

In this experimental study, panel flutter induced by an impinging oblique shockwave is investigated at a freestream Mach number of 2, using the combination of planar particle image velocimetry (PIV) and stereographic digital image correlation (DIC) to obtain simultaneous full-field structural displacement and flow velocity measurements. High-speed cameras are employed to obtain a time-resolved description of the panel motion and the shockwave-boundary layer interaction (SWBLI). In order to prevent interference between the PIV and DIC systems, an optical isolation is implemented using fluorescent paint, dedicated light sources, and camera lens filters. The effect of the panel motion on the SWBLI behavior is assessed, by comparing it with the SWBLI on a rigid wall. The results show that panel oscillations occur with a maximum amplitude of ten times the panel thickness. The dominant frequencies observed in the panel oscillation (424 Hz and 1354 Hz) match the main spectral content of the reflected shockwave position. A further POD analysis of the panel displacement spatial distribution shows that these two frequency contributions are well captured by the first two POD modes, which correspond, respectively, to a first and a third bending mode shape and account for 92% of the total oscillation energy. The fluid-structure coupling is studied by identifying, in the flow, the regions of maximum correlation between the panel displacement and the flow velocity fluctuations. The results obtained prove that the inviscid flow region upstream of the SWBLI is perfectly in phase with the panel oscillation, while the downstream region has a delay of one quarter of the flutter cycle.

@en

This experimental study investigates the use of shock control bumps (SCBs) for controlling transonic buffet. Threedimensional SCBs have been applied on the suction side of an OAT15A supercritical airfoil with the experiments conducted in the transonic–supersonic wind tunnel of Delft University of Technology at fully developed buffet conditions (Ma 0.7, α 3.5 deg and Re 2.6 × 10⁶). The effectiveness of the SCBs for different spanwise array spacings (ranging from 20 to 30%c) was verified using two optical techniques: schlieren visualization and particle image velocimetry. Both techniques confirmed the potential of controlling buffet using such devices, resulting in a reduction of the flow unsteadiness in terms of both shock oscillation and pulsation of the separated area. A dedicated particle image velocimetry investigation in a spanwise–chordwise measurement plane was conducted in order to characterize the effect of the spatial distribution of the bumps, focusing on the interaction of the shock-wave structures along the span. The configuration with a spacing of Δy_SCB 25%c was demonstrated to be the most efficient in reducing the transonic buffet oscillations and was able to reduce the reverse flow region size as compared to the clean configuration.

@en

This study experimentally investigates the effects of the sweep angle and finite wing on transonic buffet, studying two-dimensional (2-D) and three-dimensional wing configurations. Background-oriented schlieren and stereographic particle image velocimetry (PIV) have been used as measurement techniques, performing experiments on an OAT15A airfoil (clamped to both the side windows of the wind tunnel), an unswept wing, and two swept wings with sweep angles of 15 and 30 deg, respectively. The three wings are also based on the OAT15A airfoil and are clamped at the wind tunnel only at their root (free wingtip). All wings have been tested at a constant normal Mach number (Ma∞n=0.7) with respect to the leading edge. The results show that the buffet oscillations are much stronger for the airfoil than for the three finite-span wings. A large difference in the buffet behavior can be noticed between the airfoil and the unswept wing, as is also seen in oil flow visualizations. This difference is particularly evident in correspondence of the more outboard spanwise locations, suggesting that for the unswept wing, an important role could be played by finite-wing effects: notably, the tip vortex. A spectral analysis has shown that for the swept wings, the classical 2-D buffet peak (occurring at f=160  Hz for the present conditions) is substantially attenuated, whereas additional contributions in the range of 450–850 Hz appear. The PIV results showed, for the 30 deg sweep angle wing, a periodical occurrence of a secondary supersonic area downstream of the main shock wave structure, which is absent for the other wing models. The stereographic PIV configuration allowed the reconstruction of the spanwise-oriented velocity component, obtaining, in proximity of the trailing-edge, values of the spanwise velocity component (80–100  m/s) which are in agreement with the spanwise convection of buffet cells observed in the literature in this region.@en