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Premultiplied wall pressure PSD of reattachment points at different spanwise positions: (a) z ¼ 15 mm, (b) z ¼ 30 mm, and (c) z ¼ 45 mm. The red dashed line indicates the peak of premultiplied PSD at each position.
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![FIG. 4. Distribution of the time-averaged [(a)-(c)] wall pressure,...](profile/Li-Xin-Liang/publication/365128507/figure/fig2/AS:11431281095613631@1667957055394/Distribution-of-the-time-averaged-a-c-wall-pressure-d-f-skin-friction_Q320.jpg)
Swept compression ramps widely exist in supersonic/hypersonic vehicles and have become a typical standard model for studying three-dimensional (3D) shock wave/turbulent boundary layer interactions (STBLIs). In this paper, we conduct a direct numerical simulation of swept compression ramp STBLI with a 34° compression angle and a 45° sweep angle at M...
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... . (ξ, ϵ, ζ) NS OpenCFD-SCU [24,25] OpenCFD-SCU OpenCFD-SC GPU -MPI CPU OpenCFD-SC OpenCFD-SCU 200 [26] OpenCFD-SCU, 24] θ r = 120 • [24] 5 ∆x + × ∆y + × ∆z + ...
... Interactions between shock waves and turbulent boundary layers (TBLs) commonly occur in many tools of high-speed flight vehicles such as air intakes, control surfaces, and over-expanded nozzles. 1 The shock wave/turbulent boundary layer interaction (STBLI) flows have been studied in many simplified configurations, such as flow over a compression-ramp, oblique shock impinging on a flat plate with a TBL, and sharp-fin-induced shock interacting with a TBL. [2][3][4][5][6][7][8][9][10] If the interaction is strong enough, the flow is characterized by separation, which is normally associated with low-frequency unsteadiness. Such unsteady motions can cause adverse structural responses, e.g., severe buffeting of the aircraft structures, which results in structural failure and possible damage to the payload. ...
Supersonic turbulent flow over a compression ramp is studied using wall-resolved large eddy simulation with a freestream Mach number of 2.95 and a Reynolds number [based on δ0: the thickness of incoming turbulent boundary layer (TBL)] of 63 560. The unsteady dynamics of the present shock wave/turbulent boundary layer interaction (STBLI) flow are investigated by using dynamic mode decomposition techniques, linear and nonlinear disambiguation optimization, local stability analysis (LSA), and global stability analysis (GSA). By analyzing the dynamic system for the STBLI flow, three dynamically important modes with characteristic spanwise wavelengths of 2δ0, 3δ0, and 6δ0 are captured. The 2δ0 mode approximates the spanwise scale of the Görtler-like vortices and Görtler mode of LSA, suggesting the presence of Görtler instability, which is believed to be related to the unsteady motion of streaks downstream of reattachment in the flow. The features of the 3δ0 mode are also observed in large-scale motions of the incoming TBL, implying the existence of a convective mechanism that is excited and maintained by such motions. Additionally, the GSA results show the most unstable mode features a spanwise wavelength of around 6δ0, indicating the existence of global instability that is believed to be related to the oscillating motion of separation shock. The coexistence of these three mechanisms is confirmed. Discussions on the above findings provide an interpretation for low-frequency unsteadiness that the unsteadiness of surface streaks results from the combined effects of the Görtler instability near flow reattachment and the convection of large-scale motions in the incoming boundary layer, while the low-frequency shock motion may be related to a global mode driven by upstream disturbances.
... As such, several academic configurations have been recognised and discussed in literature. These include (a) normal shock interactions (Atkin & Squire, 1992;Bruce & Babinsky, 2008;Carroll & Dutton, 1990), (b) incident-reflecting shock interactions (Baidya et al., 2020;Dupont et al., 2006;Humble et al., 2009b;Pasquariello et al., 2017;Piponniau et al., 2009;Souverein et al., 2013;Touber & Sandham, 2011), (c) blunt fin interactions (Brusniak & Dolling, 1994;Dolling & Bogdonoff, 1982;Hou et al., 2004;Hung & Buning, 1985;Ngoh & Poggie, 2022), (d) two-dimensional (Dolling & Murphy, 1983;Dolling & Or, 1985;Erengil & Dolling, 1991;Ganapathisubramani et al., 2009;Loginov et al., 2006;Schreyer et al., 2018;Settles et al., 1979;Smits & Muck, 1987;Sun et al., 2020;Wu & Martín, 2008) and three-dimensional (Adler & Gaitonde, 2019;Ceci et al., 2023;Erengil & Dolling, 1993;Settles et al., 1980;Vanstone et al., 2018;Zhang et al., 2022) compression-ramp interactions, etc. ...
Shock-wave/boundary-layer interactions (SWBLIs) are complex flowfields commonly encountered in many high-speed aerospace applications. A strong shock wave imposes a large adverse pressure gradient and can lead to large-scale flow separation. The separated flow and the separation-inducing shock are highly unsteady, causing fluctuating pressure and thermal loads on the surface, inlet instabilities, et cetera; they are generally detrimental to the safe and reliable operation of the vehicle. Hence proper control of the flow is essential. In this thesis, separation control with air-jet vortex generators (AJVGs) is experimentally investigated. Various configurations of a single row of spanwise-inclined AJVGs at 45° are used to control a fully-separated 24° compression-ramp interaction. The results show clear evidence of streamwise vortices induced by the injected jets. Consequently, the 2D separation line of the baseline case is now corrugated. Effectively configured AJVGs can reduce separation length by nearly 25%, the recirculation region area by as much as ~ 57%, and alleviate turbulence amplification across the SWBLI. However, these effects strongly depend on various AJVG control parameters, amongst them are jet-to-jet spacing, injection pressure, orifice shape, etc. Investigations of jet/jet spacing show that the jet-induced streamwise vortices incite different control effects based on the degree of interaction between adjacent vortices; AJVGs with intermediate spacing display the most favourable control effects due to the formation of stable, interacting and streamwise-elongated coherent vortical structures downstream of the jet array. A similar mechanism occurs when varying the air-jet injection pressure, where intermediate injection pressures show the best control effectiveness, while both very large and very small injection pressures are ineffective. Furthermore, the results highlight a strong interplay between jet spacing and injection pressure in defining the flow control effectiveness of AJVGs. A more favourable control effect can also be brought about by using non-circular AJVGs, and in-particular, elliptical AJVGs. Measurements of AJVGs with elliptical orifices reveal their improved control effectiveness due to better flow entrainment and higher turbulent mixing, compared to equivalent circular AJVGs. This is attributed to stronger streamwise vortices that penetrate, on average, 25% farther into the boundary layer. For all AJVG-controlled cases, the separation bubble and shock exhibit the characteristic low-frequency unsteadiness observed for the baseline case. However, the dynamics of unsteadiness remain largely unaltered by AJVGs. Conditional analysis of the PIV results show that the separated shear-layer dynamics and the separation-bubble motion are dominant, despite the introduced coherent structures and the modification of the incoming boundary layer due to jet-array injection.
... As a consequence, such phenomena and interactions have been-and still are-investigated, both experimentally 16,17,[22][23][24][25][26][27] and numerically. [18][19][20][28][29][30][31][32][33][34][35][36][37][38][39][40][41][42][43][44][45][46] Provided that the shock wave is sufficiently strong, a recirculation bubble in the boundary layer forms where the shock impinges on the surface of the aero-structure. An example of such a shock-bubble interaction region is illustrated in Fig. 2. Observations indicate that the shock-bubble interaction can oscillate with a characteristic frequency of one or two orders of magnitude lower than the typical frequency of the supersonic turbulent boundary layer (TBL) and with a length scale much larger than the TBL thickness. ...
This paper investigates deep learning methods in the framework of convolutional neural networks for reconstructing compressible turbulent flow fields. The aim is to develop methods capable of up-scaling coarse turbulent data into fine-resolution images. The method is based on a parallel computational framework that accepts five image sets of various resolutions, trained to correspond to the respective fine resolution. The network architecture mainly consists of convolutional layers, constructing an encoder/decoder network. Based on the U-Net scheme, three different implementations are presented, with residual and skip connections. The methods are implemented in a supersonic shock-boundary-layer interaction problem. The results suggest that simple networks perform better when trained on limited data, and this can be a practical and fast solution when dealing with turbulent flow data, where the computational burden is most of the time difficult to decrease. In such a way, a coarse simulation grid can be upscaled to a fine grid.
... Three-dimensional effects are considered by Di Renzo et al. 47 , Larsson et al. 48 , Zhang et al. 49 . In recent studies, Li et al. ...
In the present study, we investigate the evolution of turbulent statistics and coherent structures in hypersonic turbulent boundary layers at the Mach number of 5 impinged by oblique shock waves generated by the wedge with the angles of 14°, 10°, and 6°, inducing strong, mild, and incipient flow separation, by exploiting direct numerical simulation databases, for the purpose of revealing the underlying flow physics that are of significance to turbulent modeling. We found that the large-scale structures are amplified within the interaction zone, manifested in the form of large-scale low- and high-speed streaks with the spanwise length scale of boundary layer thickness, and gradually decay downstream, the process of which is extremely long. The abrupt variation in the characteristic length, time, and velocity scales as well as the incompatible viscous dissipation of the mean and turbulent kinetic energy results in the incorrect predictions by the Reynolds-Averaged Navier–Stokes (RANS) equation simulations, provided the models are established based on solving the transport equations of the turbulent kinetic equation and its viscous dissipation (k−ε or k−ω models, for instance). To amend this issue, we propose to refine the parameters in the model as the functions of wall pressure, the flow quantities related to multiple flow features. The RANS simulations with the k−ω SST model utilizing the proposed refinement improve greatly the accuracy of the skin friction, wall heat flux, and Reynolds shear stress downstream of the interaction zone, and the wall pressure distributions in hypersonic turbulence over compression ramp, suggesting its promising prospect in engineering applications.
... Xie et al. 17 studied the characteristics and its variation with Reynolds number in expansioncompression corner. Zhang et al. 18 conducted a DNS of swept compression ramp STBLI. They found that the spanwise variation of the shock wave inclination induced the peak friction and heating increase along the spanwise direction. ...
Large eddy simulations of shock wave/turbulent boundary layer interaction on a compression ramp at the Mach number [Formula: see text] and Reynolds number [Formula: see text] are performed to investigate the impact of the incipient and fully separated conditions on the development of the flow field. The quasi-dynamic subgrid-scale kinetic energy equation model, which combines the benefits of the gradient model with the eddy-viscosity model, has been applied. Compared with the previous experimental and numerical results, the simulation was validated. The flow structures, turbulence properties, vortex structures, and low-frequency unsteadiness are all investigated. The flow field of the incipient separation is attached and rarely impacted by shock. An evident separation bubble and localized high wall temperatures in fully separated flow are caused by the separation shock's significant reverse pressure gradient. The Reynolds stress components exhibit significant amplification in both cases, and the peak outward shifts from the near-wall region to the center of the free shear layer. Turbulent kinetic energy terms were analyzed, and the two scenarios show a significant difference. The power spectral density of the wall pressure fluctuations shows that the low-frequency motion of the incipient separation is not apparent relative to the fully separated flow.
... However, if flow conditions are globally unstable and perturbations are introduced, the dynamcis of the thermal streaks are still not fully understood, and it is likely to be determined by both convective and intrinsic mechanisms. Most shock-wave/turbulent boundarylayer (STBLI) interactions with large separation belong to this category [44][45][46][47][48][49][50] , as the turbulent boundary layer continuously introduces perturbations into the separated flow. The convective/intrinsic mechanisms may also be a potential way to explain the other dynamics that occurred in STBLI flows (e.g., low-frequency motions of separation bubbles), which is very worthwhile for future study. ...
Hypersonic laminar flow over a canonical 25‒55{degree sign} double cone is studied using computational fluid dynamics, bispectrum analysis, and dynamic mode decomposition (DMD) with a freestream Mach number of 11.5 and unit Reynolds number of 1.6 × 10 ⁵ m ⁻¹ . The present study focuses on the evolution and nonlinear behavior of perturbation modes in the flow. The presence of the perturbation modes is first described in detail through the results of direct numerical simulation (DNS). The results of high-order spectrum analysis (bispectrum) then reveal complex nonlinear interactions in the flow. By examining the evolution of such interactions, the frequency broadening phenomenon of the fully saturated flow is explained, and the unsteady dynamics of the fully saturated flow are recognized to be caused by the nonlinear saturation of linear instability in the flow. This causality is further confirmed by the DMD results of the Stanton number near the reattachment region. The origins and dynamics of unsteady saturated flow in the hypersonic laminar flow are therefore demonstrated.
We carry out a numerical study of swept shock wave/turbulent boundary layer interaction in the hypersonic regime. Starting from a numerical/experimental benchmark case of a nearly adiabatic two-dimensional hypersonic interaction, a crossflow velocity component is added to the incoming flow to mimic three-dimensional interactions with cylindrical symmetry. We observe, for a fixed streamwise Mach number, monotonic increase of the extent of the interaction region for the swept cases. An attempt at extending the free-interaction theory to hypersonic swept interactions is made, which is found to apply only to the initial part of the interaction region. The spatiotemporal dynamics of wall pressure on mean separation line features large-scale pressure corrugations, which are advected at a phase speed which is a fraction of the mean crossflow velocity, if present. The characteristic wavelength of the corrugation is found to be a multiple of the separation bubble size. The numerically estimated peak frequencies well conform with the previously introduced formula for swept supersonic interactions [Ceci et al., J. Fluid Mech. 956, R1 (2023)]. Proper orthogonal decomposition is applied to investigate the spatial structure of the corrugation at the separation point and educe phase relations between the flow structure and pressure oscillations at the reattachment point. The present analysis leads us to conclude that the same phenomenology found in swept supersonic interactions also holds in the hypersonic case.
This paper introduces open-source computational fluid dynamics software named open computational fluid dynamic code for scientific computation with graphics processing unit (GPU) system (OpenCFD-SCU), developed by the authors for direct numerical simulation (DNS) of compressible wall-bounded turbulence. This software is based on the finite difference method and is accelerated by the use of a GPU, which provides an acceleration by a factor of more than 200 compared with central processing unit (CPU) software based on the same algorithm and number of message passing interface (MPI) processes, and the running speed of OpenCFD-SCU with just 512 GPUs exceed that of CPU software with 130000 CPUs. GPU-Stream technology is used to implement overlap of computing and communication, achieving 98.7% parallel weak scalability with 24576 GPUs. The software includes a variety of high-precision finite difference schemes, and supports a hybrid finite difference scheme, enabling it to provide both robustness and high precision when simulating complex supersonic and hypersonic flows. When used with the wide range of supercomputers currently available, the software should able to improve the performance of large-scale simulations by up to two orders on the computational scale. Then, OpenCFD-SCU is applied to a validation and verification case of a Mach 2.9 compression ramp with mesh numbers up to 31.2 billion.
