Shock Wave Boundary Layer Interaction Research Articles

A universal scaling for the length scales of shock-induced separation

Experiments of transitional shock wave–boundary layer interactions (SBLIs) over 6 $^\circ$ and 10 $^\circ$ compression ramps were performed at Mach number 1.65. The unit Reynolds number was varied by a factor of two between 5.6 million per metre and 11 million per metre. Schlieren flow visualization was performed, and mean flow measurements were made using Pitot probes. Free interaction theory was verified from pressure measurements for all operating conditions. A new non-dimensional parameter was developed for scaling the strength of the imposed shock, which was based on the pressure required to separate a boundary layer. The validity of this new scaling was supported by the reconciliation of large discrepancies in a diverse collection of experimental results on the length scales of transitional interactions. This non-dimensional scaling was also applied to turbulent interactions, where different models were used to determine the pressure required to separate a turbulent boundary layer. Finally, a direct comparison between transitional and turbulent SBLIs was made, which revealed new insights into the evolution of length scales based on the state of the boundary layer.

Spectral proper orthogonal decomposition of external flow at high Reynolds number

Spectral proper orthogonal decomposition(SPOD) has been recently used to capture temporal and spatial characteristics of complex flows. Applying SPOD analysis to such flow fields can deepen the understanding of flow mechanism. To further study the external flow at high Reynolds numbers, large eddy simulation(LES) is applied to acquire the flow field data while SPOD is utilized to extract important flow structures. The considered external flow includes incompressible flow around the cylinder and transonic shock oscillation. For the flow around the cylinder, flow field under three different Reynolds numbers is studied. Power spectral density(PSD) results of lift coefficient show that the vortex shedding frequency of low Reynolds number is St=0.24, which of moderate Reynolds number is St=0.203, and St=0.27 for high Reynolds number flow. The increment of Reynolds number will lead to early instability of shear layer on the cylindrical surface. The modal results of SPOD and dynamic mode decomposition(DMD) at different Reynolds numbers show two antisymmetric structures at shedding vortex frequencies, indicating the resemblance in the process of vortex formation. For the shock wave/boundary layer interaction(SBLI) of transonic Royal Aircraft Establishment(RAE) 2822 airfoil, flow field under M∞=0.729, Rec=1.94×107 is studied. The results show that the occurrence of unsteadiness is closely related to the normal shock wave/turbulent boundary layer interaction happening on the upper surface of the airfoil. Modal results of SPOD near the characteristic frequency show strong discontinuity near the shock wave.

High-bandwidth pressure field imaging of fin-generated shock wave–boundary layer interactions

The dynamics of a shock-induced separation unit generated by a 20 $^\circ$ sharp fin placed on a cylindrical surface in a Mach 2.5 flow was investigated. Specifically, the present work investigated the mechanisms that govern the mid-frequency range of separation shock unsteadiness in the fin shock wave–boundary layer interaction (SBLI) unit. Two-dimensional pressure fields were obtained over the cylinder surface spanning the entire fin SBLI unit using high-bandwidth pressure-sensitive paint at 40 kHz imaging rate that allowed probing the low- through mid-frequency ranges of the separation shock unsteadiness. The mean pressure field showed a progressive weakening of the separation shock with downstream distance, which is an artifact of the three-dimensional relief offered by the curved mounting surface. The root-mean-square (r.m.s.) pressure field exhibited a banded structure with elevated $p_{r.m.s.}$ levels beneath the intermittent region, separation vortex and adjacent to the fin root. The power spectral density (PSD) of the surface pressure fluctuations obtained beneath the intermittent region revealed that the separation shock oscillations exhibited the mid-frequency content over the majority of its length. Interestingly, neither the PSD nor the length of the intermittent region varied noticeably with downstream distance, revealing a constant separation shock foot velocity along the entire SBLI. The pressure fluctuation PSD beneath the separation vortex also exhibited the broadband peak at the mid-frequency range of the separation shock motions over the majority of its length within the measurement domain. By contrast, the region adjacent to the fin root exhibited pressure oscillations at a substantially lower frequency compared with the separation shock and the separation vortex. Two-point coherence and cross-correlation analysis provided unique insights into the critical sources and mechanisms that drive the separation shock unsteadiness. The separation vortex and separation shock dynamics were found to be driven by a combination of convecting perturbations that originated from the vicinity of the fin leading edge and the local interactions of the separated flow with the incoming boundary layer. The boundary layer locally strengthened or weakened the convecting pressure perturbations depending on the local momentum fluctuations within the boundary layer.

Aerodynamic heating in hypersonic shock wave and turbulent boundary layer interaction

In hypersonic flight the shock wave and turbulent boundary layer interaction (STBLI) sharply increases wall heat transfer that intensifies the aerodynamic heating problems. In this work the STBLI is modelled by compression ramp flow with a Mach number of 5, a Reynolds number based on momentum thickness of 4652 and a wall to recovery temperature ratio of 0.5. The aerodynamic heat generation and transport mechanisms are investigated in the interaction based on theoretical analysis and direct numerical simulation (DNS) that agrees with previous studies. A prediction correlation of wall heat flux in STBLI is deduced theoretically and validated by some representative data including the present DNS, which improves the prediction accuracy and can be applied to a wider $Ma$ range compared with the canonical Q-P theory. The correlation indicates that the sharp increase of wall heat transfer in the STBLI can be explained by the boundary layer compression and the convection transport enhancement. Based on the DNS results, the aerodynamic heat generation and transport mechanisms are revealed in the separation, recirculation and reattachment zones in the STBLI. From this perspective, the peak heat flux can be further explained by the enhancement of near-wall turbulent energy dissipation, compression aerodynamic heat generation and the near-wall turbulent transport. The generation and transport of compression aerodynamic heat reveal the underlying mechanism of the strong correlation between the peak heat flux ratios and the pressure ratios in STBLIs.

Evaluation of passive control systems for shock wave boundary layer interaction within a supersonic air inlet

Evaluation of passive control systems for shock wave boundary layer interaction within a supersonic air inlet

Highly separated transitional shock-wave/boundary-layer interactions: A spatial modal study

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.

Shock Oscillation Mechanism of Highly Separated Transitional Shock-Wave/Boundary-Layer Interactions

Reynolds numbers at cruise altitude can be such that a laminar boundary layer persists on the suction side of a transonic fan blade up to the shock-wave/boundary-layer interaction (SBLI). In a transitional SBLI which exhibits sufficiently large shock-induced separation, a shock oscillation mechanism characterized by growth and natural suppression of the upstream laminar section of the separation bubble occurs. To validate the shock oscillation mechanism observed in large eddy simulations (LES), the shock oscillation mechanism is studied experimentally using high-speed Schlieren and spark-light shadowgraphy. A characteristic length based on the distance of laminar separation shock travel is proposed. Strouhal numbers from LES and the experiment collapse at around 0.075. A strong dependency of the oscillation mechanism on free-stream turbulence and boundary-layer state is shown. Dominant oscillation frequencies are an order of magnitude lower for the turbulent interaction as opposed to the laminar case. For the laminar case, dynamic mode decomposition showed a strong relationship of the laminar separation shock with the separation bubble and reflected shock movement. The turbulent interaction shows a significantly lower reflected shock travel distance. The findings experimentally confirm that stabilization of the shock is achieved by tripping the boundary layer.

Shock waves characteristics and losses estimation of non-equilibrium condensation flow in nozzle and steam turbine cascade

The investigation of transonic non-equilibrium condensation is paramount due to its profound influence on the shock patterns and the distribution of losses. This study numerically investigates the shock wave morphology, evolution, and quantitative loss calculation under the influence of non-equilibrium condensation within the Moses-Stein nozzle and Dykas cascade. The results indicate that the vapor near the nozzle wall undergoes phase transition first and forms “X-shaped” condensation shock, which impervious to back pressure variations. Conversely, the “λ-shaped” aerodynamic shock moves towards the throat with increasing back pressure, triggering boundary layer separation. The entropy generation of “λ-shocks” accounts for over 98% and represents the primary source of shock losses. Strong pressure side shock induces shock wave-boundary layer interactions with the blade suction wall, resulting in reflected shocks. However, the latent heat released by non-equilibrium condensation mitigates this effect, increasing the total pressure loss coefficient by 0.188. The decrease in Mach number under high back pressure causes a notable increase in the trailing edge shock angle, with the wave foot shifting upstream. Notably, the proportion of shock losses attributed to the suction side of trailing edge exceeds 80%, significantly surpassing other regions, which progressively increases with rising back pressure. Therefore, it is crucial for cascade optimization. The findings can provide references for the optimization design of low-pressure steam turbine blade profiles. The results can serve as a guide for optimizing the blade profiles of low-pressure steam turbine.

Study on the flow characteristics of double-cone in hypersonic flows

Reusable spacecraft is one of the hot topics in the field of aerospace, attracting wide attention from researchers. Due to the high-Mach-number flight of re-entry, the air near the wall exhibits significant thermochemical nonequilibrium effects. Besides, the existing shock wave/boundary layer interaction (SWBLI) leads to severe aerodynamic heating issues. This study utilizes the two-temperature model to conduct simulations of hypersonic laminar flows around a canonical 25°/55° double-cone with low and high enthalpy. By varying the wall temperature and the aft cone angle, the evolution mechanism of the flow under the high-enthalpy and hypersonic freestream is explored. Our findings illustrate that while the thermodynamic nonequilibrium model closely reflects the separation zone evidenced in the experiments, it habitually overpredicts the peak heat transfer, particularly under high-enthalpy conditions. For the thermochemical nonequilibrium model, the flow field structure appears more uniform, with a reduced standoff distance of the detached shock. An incremental rise in wall temperature correlates with a proportional augmentation of the separation bubble, though its impact on the overall flow field is negligible. Increasing the aft cone angle intensifies the shock/shock interaction (SSI), transitioning from a lower-intensity Type VI to a more intense Type IV shock interplay. The examination reveals that the increase in temperature and cone angle amplifies the interaction between the separation region and shock waves, drastically escalating the peak heat transfer and fostering a secondary peak. The hypersonic flow of the double-cone demonstrates a multifaceted interaction of phenomena, notably SWBLI and SSI, with our analyses providing pivotal insights for the aerodynamic and thermal protection design of the high-Mach-number spacecraft.

Effect of multi-temperature models on heat transfer and electron behavior in hypersonic flows

In hypersonic computational fluid dynamics, the two-temperature (2-T) model is widely used to simulate thermochemical nonequilibrium. The 2-T model incorporates translational-rotational and electron-electronic-vibrational energies, assuming that the integrated energies have the equivalent temperature. In this study, multi-T models are constructed to accurately predict the effects on heat flux and free electrons due to the separation of energy modes under hypersonic environments. The three-temperature (3-T) model separates the electron-electronic energy from the electron-electronic-vibrational energy of the 2-T model. The 3-T model can accurately predict the distribution and temperature of free electrons by separating the energy of free electrons, which has different characteristics from heavy particles. The four-temperature model treats rotational energy as a nonequilibrium energy mode, distinct from translational-rotational energy. While the rotational temperature reaches equilibrium rapidly at low temperatures, at high-temperature regime rotational temperature shows a relaxation time similar to that of vibrational temperature, which cannot be ignored. To develop multi-T models, electron-vibrational relaxation and translational-rotational relaxation, which are omitted in the 2-T model, are considered. Various flight test and ground facility conditions are analyzed to verify the effects of electron and heat flux under circumstances that include shock, expansion, and shock wave boundary layer interaction. The results of the multi-T models show significant differences in electron temperature and distribution caused by electron-electronic nonequilibrium. Additionally, rotational nonequilibrium increases the shock standoff distance and alters the electron distribution at high altitudes. The heat flux difference across multi-T models is found to be negligible, except in the high degree of ionization condition.

Pressure feedback system for flow separation mitigation in scramjet intakes

Shock Wave Boundary Layer Interactions (SWBLI) occur due to the convergence of a shock wave and a viscous boundary layer. The resulting separation bubble is a major setback for the performance of high-speed air intakes. This paper presents a numerical investigation into the potential of a Pressure Feedback Technique (PFT) to alleviate shock-induced flow separation within a Scramjet engine intake operating at Mach 4. The PFT is a novel self-sustaining flow control technique that combines simultaneous suction and injection. The two main output parameters used to support the hypothesis of the present study are the size of the separation bubble as well as the total pressure recovery at the isolator outlet. The most prominent observation of this study is that with the installation of the PF tubes, the total pressure recovery at the exit is noted to rise. The effectiveness of the pressure feedback technique in reducing the intensity of the separation bubbles is found to depend on the diameter of the PFT, pitch-to-diameter ratio, and PFT tube design.

Extremely high wall pressure events in shock wave and turbulent boundary layer interactions using DNS data

This study investigates high-amplitude Extreme Wall Pressure fluctuation Events (EWPEs) in Shock wave/Turbulent Boundary Layer Interactions (STBLIs) through the conditional sampling of direct numerical simulation databases. The aim is to evaluate the effect of STBLIs and their strength on the statistical properties and associated turbulent structures of EWPEs using the conditional-averaging and clustering method. The temporal statistical results show that the occurrence probability and contribution ratio of EWPEs decrease downstream of strong STBLI, but their duration and interval time increase. Regarding two-dimensional wall pressure structures, the large population of small-scale structures becomes more elongated, but strong interactions induce a greater number of large-scale structures. The pairing of wall pressure events with a higher occurrence probability is verified by the joint probability density functions. Conditional analysis reveals that, as the interaction strength increases, the ejection motion associated with positive events occurs farther downstream and the spanwise vortex core locating above negative events is lifted up along the wall-normal direction. Moreover, analysis associates the paired wall pressure events with the sweep, ejection, and swirl motions in STBLIs, where hairpin eddies play an important role in the formation of positive–negative paired wall pressure structures.

Large eddy simulation of shock wave/boundary layer interactions in a transonic compressor cascade

In this paper, large eddy simulation (LES) was performed to investigate the shock wave/boundary layer interaction (SBLI) phenomenon in transonic compressor cascades with a chord Reynolds number of 2.12 × 106. A comprehensive analysis was conducted on both the SBLI structures inherent to the transonic compressor cascade and the coherent vortex structures within the boundary layer. The underlying mechanisms of the shock-induced boundary layer transition and the shock low-frequency unsteadiness in the transonic compressor cascade were elucidated through spectral and dynamic mode decomposition (DMD) analysis. The results revealed that boundary layer separation induced by the SBLI cannot reattach, leading to the formation of large-scale coherent vortex structures. Spectral analysis revealed that the shock-induced boundary layer transition in the transonic compressor cascade was dominated by inviscid Kelvin–Helmholtz (K–H) and secondary instability mechanisms, characterized by a dimensionless Strouhal number of 0.06. Additionally, pressure signals showed the variations in sub-frequency from the separated shear layer to the main flow. The oscillation amplitude of the shock foot was significantly greater than that of the shock main body, and the oscillation frequency of the shock foot was consistent with the sub-frequency. The oscillation frequency of the shock main body coincided with that of the compression ramp and flat plate configurations. Finally, DMD modal analysis indicated that high-frequency modes were correlated with turbulent fluctuations in the boundary layer, while medium- and low-frequency modes corresponded to shedding motion in the separated shear layer and low-frequency motion of the shock. This work promotes the understanding of the complex flow mechanisms of SBLI in the transonic compressor cascades.

Sidewall influence of varying free stream Mach number in ramp induced shock wave boundary layer interactions

This study investigates the three-dimensional (3D) effects introduced by the end walls for an aspect ratio of 1 in ramp-induced shock wave boundary layer interactions. The simulations are performed using a symmetry boundary condition in the spanwise direction at free-stream Mach numbers in 3D. The simulations are performed using an in-house compressible supersonic solver “OpenSBLIFVM”. Two free stream Mach numbers 2.5, and 3 are used in the current work, and the simulated results are compared with the aspect ratio 1 simulations by Mangalagiri and Jammy. The inflow is initialized with a similarity solution; its Reynolds number based on the boundary layer thickness is adjusted such that the Reynolds number at the start of the ramp is kept at 3×105 for all simulations. From the results, it is evident that the introduction of sidewalls resulted in a shorter centerline separation length when compared with the two-dimensional (2D) simulations. This contradicts the results at Mach 2 by Mangalgiri and Jammy where the vortex observed at Mach 2 in the central separation region disappeared with increasing free-stream Mach number. Additionally, the topology of interaction shifted from owl-like separation of the second kind to the first kind when the freestream Mach number increased from 2 to 2.5. It can be concluded that the interaction topology is crucial to the increase or decrease of the central separation length when compared to 2D simulations.

Assimilating mean velocity fields of a shockwave–boundary layer interaction from background-oriented schlieren measurements using physics-informed neural networks

We propose a novel method to reconstruct mean velocity fields of turbulent shockwave–boundary layer interactions (SBLIs) from background-oriented schlieren (BOS) measurement data using physics-informed neural networks (PINNs). By embedding the compressible Reynolds-Averaged Navier–Stokes equations into the PINN loss function, we recover a full set of physical variables from only the density gradient as training data. This technique has the potential to generate velocity fields similar to particle image velocimetry (PIV) results from usually simpler planar BOS measurements, at the cost of some computational resources. We analyze our method's capabilities on two oblique SBLI cases: a high-fidelity Mach 2.28 direct numerical simulation dataset for validation and a Mach 2.0 wind tunnel experiment. We demonstrate the positive impact of different wall boundary constraints such as the wall shear stress and pressure distribution for enhancing the PINN's convergence toward physically accurate solutions. The predicted fields are compared with experimental PIV and other point measurements, while we discuss the accuracy, limitations, and broader implications of our approach for SBLI research.

Dynamics of a 3-D inlet/isolator measured with fast pressure-sensitive paint

Fast pressure-sensitive paint (PSP) was applied to an inlet/isolator designed using the Osculating Internal Waverider Inlet with Parallel Streamlines (OIWPS) method. The dorsal isolator surface pressure was measured using anodized-aluminum PSP through transparent cast acrylic that makes up the ventral portion of the isolator. Temperature-sensitive paint was utilized to correct for the PSP’s temperature sensitivity. The model was tested under Mach 5.7 flow at Re =\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$=$$\\end{document} 8.5 ×106\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\ imes 10^6$$\\end{document} /m and 10.2 ×106\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\ imes 10^6$$\\end{document} /m in the AFOSR–Notre Dame Large Mach-6 Quiet Tunnel (ANDLM6QT) under conventional noise conditions. Flow phenomena, such as shocks originating in the inlet and flow separation at the throat, were visualized with high spatial resolution. The dynamics measured by the PSP and pressure transducers matched well where the spectral signal-to-noise ratio was above unity. Power spectral densities showed significant frequency content at ≈\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\\approx$$\\end{document}1 kHz in the shock-wave/boundary-layer interaction (SWBLI) regions. Coherence analysis showed a linear relationship between the unsteady pressures at locations underneath different SWBLI in the isolator, with the exception of the Busemann throat shock. Temporal correlation of shock positions indicated that disturbances propagated downstream at 114% of the core-flow velocity; however, improved calculations of the core-flow velocity are needed to refine this assessment. The surface pressure fields at Re = 8.5 ×106\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\ imes 10^6$$\\end{document} /m and 10.2 ×106\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\ imes 10^6$$\\end{document} /m were quantitatively very similar, and the results in the ANDLM6QT were qualitatively similar to previous studies in the Boeing/AFOSR Mach-6 Quiet Tunnel under noisy flow.

Alleviation of SWBLI over the payload of a launch vehicle by change of nose shape

Wind tunnel tests on a typical heavy-lift launch vehicle configuration featuring a core vehicle flanked by two strap-on boosters showed very high levels of unsteady pressures over the payload region at transonic Mach numbers when the nose had a cone-cylinder shape. This was due to very high levels of Shock Wave Boundary Layer Interactions (SWBLI) associated with the conical shape, which caused alternating flow featuring the formation of a unsteady λ-shock system over the payload region and a pair of counter-rotating vortices. The phenomenon was found to strongly depend on the proximity of the noses of the strap-on boosters to the boat tail downstream of the payload region. When the conical nose was replaced by an ogive-shaped nose, significant reduction in SWBLI occurred, with dramatic reduction of shock oscillations, pressure fluctuations, replacement of alternating flow by steady flow and disappearance of the vortex-pair. Time-averaged pressure fluctuations indicate reduction in pressure gradients due to a more gradual expansion of flow over the ogive nose compared to conical nose, an observation also supported by CFD simulations. The alleviation of SWBLI with ogive nose was so effective that the presence or absence of strap-on boosters had practically no effect on pressure fluctuations over the payload region. Unlike the conical shape, the ogive nose cone shape seems to meet both the shape and pressure parameter-criteria of NASA to ensure buffet-free environment.

Experimental study on the hypersonic double incident shock wave/boundary layer interaction regulated by plasma actuation array

In the inlet passage of a hypersonic vehicle, multi-channel shock waves inevitably interact with the boundary layer, producing complex multi-channel shock wave/boundary layer interactions (SWBLIs). The flow separation caused by these interactions significantly decreases the intake efficiency and may prevent the intake from starting. The typical interaction mode of multi-channel interactions is through double incident SWBLIs. Therefore, it is necessary to study the characteristics of double incident SWBLIs and identify relevant flow control techniques. In this paper, the characteristics of hypersonic double incident SWBLIs are first examined, and then the results of an experimental study on regulation using a plasma actuation array are reported. We find that plasma actuation can positively regulate the hypersonic double incident SWBLI, and the optimal control effect reduces the area of the separation bubble by 38.62%. The main regulation mechanism involves suppressing the low-frequency instability of SWBLIs through a high-frequency shock effect. The regional scale of the separation bubble can be controlled by regulating the shock wave oscillation range. Correlative results provide technical and method support for the application of plasma actuation in hypersonic double incident SWBLI regulation and present a new idea for the selection of flow control methods for advanced intake systems.

Spanwise unsteadiness in the sidewall-confined shock-wave/boundary-layer interaction

Three-dimensional effects of sidewalls on the low-frequency unsteadiness of the shock-wave/boundary-layer interaction (SBLI) are of academic and practical importance but not yet well understood. Considerable attention has been paid to the viscous effect of sidewalls, whereas the potential inviscid confinement effect of sidewalls has received little attention. The present work provides experimental evidence of multiscale spanwise travelling waves crossing the separation front under the confinement of sidewalls. Global pressure measurements were made for a sidewall-confined 24 $^\circ$ compression ramp interaction in Mach-2.83 flow using fast-responding pressure-sensitive paint. The unsteady pressure in a statistically two-dimensional intermittent region suggests that in addition to the canonical streamwise oscillation, the separation front exhibits significant low-frequency, multiscale spanwise distortion. Modal analysis further reveals that multiscale spanwise unsteadiness has higher intensity and frequency than the streamwise oscillation. Such strong spanwise unsteadiness calls attention to the low-frequency unsteadiness in previous sidewall-confined SBLI experiments and encourages further study on the mechanism of the confinement effect.

Low-frequency unsteadiness in hypersonic swept shock wave-boundary layer interactions

Low-frequency unsteadiness in hypersonic swept shock wave-boundary layer interactions