Shock Boundary Layer Interaction Research Articles

Comparative study of transonic shock–boundary layer interactions due to surface heating and cooling on an airfoil

Implicit large eddy simulation results are compared to investigate the effects of wall-heating and wall-cooling on shock–boundary layer interaction over an airfoil. Heat flux is provided on the suction surface of the airfoil from x/c=0.40 to x/c=0.50 for a Mach number of 0.72 and a Reynolds number based on chord of Re=16.2×106. Flow quantities are compared for the effects of heating and cooling. Numerical Schlieren snapshots reveal an oscillation of the shock wave and its interaction with upstream propagating Kutta waves generated from the trailing edge of the airfoil. Quantitative data obtained from these Schlieren snapshots and the mean aerodynamic load values indicate a reduction in frequency of oscillation of shock wave and a decrease in shock strength for the case of heating. Flow control by heating shows higher fluctuations in flow features evident from instantaneous quantities. Both imposed excitations lead to a marginal increase in aerodynamic efficiency (lift/drag). We also compare the integral aerodynamic parameters, such as lift and drag coefficients, and their ratio, Cl, Cd,and Cl/Cd. The simulations reported here follow the techniques used in Sengupta et al. [“Thermal control of transonic shock–boundary layer interaction over a natural laminar flow airfoil,” Phys. Fluids 33(12), 126110 (2021)].

Wall-Modeled Large Eddy Simulation of High Speed Flows

The use of wall-modeled large-eddy simulation (WMLES) is explored in the context of compressible flows with a focus on cold-wall boundary layers and flows with shock-induced separation. It is observed that for cold-wall flows, a “mixed” scaling for the length scale appearing in the eddy viscosity formulation outperforms the classical semilocal scaling for obtaining predictions of heat flux and skin friction. A few shock/boundary-layer interaction (SBLI) cases are examined in some detail, and model modifications are proposed to overcome identified deficiencies. It is shown that using WMLES the low-frequency characteristics of SBLI at high Reynolds number can be quantitatively captured. A dynamically switched version of the equilibrium model is proposed; this shows promise for relatively inexpensive simulations at these conditions.

CFD investigation of shock boundary layer interaction in hypersonic flow and flow control using micro ramps

PurposeThis study aims to investigate the prevalence of shock boundary layer interaction (SBLI) in air-breathing intake system is highly undesirable since this leads to high pressure gradients, typical stream mutilation and pressure drop. A novel flow control mechanism is incorporated in this research holding an array configuration of passive flow control device (micro ramps [MR]) that is adapted to improve the boundary layer stability.Design/methodology/approachTwo geometric variants of the MR, namely, MR40 and MR80 is considered which reduce the pressure drop during SBLI. The incidence oblique shock wave angle of 34° is considered for the modelling. Large eddy simulation (LES) turbulence model was used with subgrid models of Wall modelled LES, Smagorinsky–Lilly to compute the unsteady effects of SBLI control using micro vortex generators. The unsteady results are compared with steady Reynold’s average Naviers–Stoke’s equation for calibrating the turbulence models.FindingsThe array configuration of MR80 reduces the pressure drop by 22% as compared with no ramp configuration and also reduces the flow distortion in hypersonic inlet. The most affected region of the MR is in the vicinity of center-line. Quantitative results prove that the upstream influence of the shock waves has been largely reduces by MR80 array configuration as compared to single MR80 pattern configuration. Different vortex structures found in the experiments was exclusively predicted using LES.Originality/valueThis paper substantiates the requirement of MR array configuration for transferring the momentum from free stream to the boundary layer and thereby energizing the boundary layer. This process of energization delays the flow separation in hypersonic flow.

Evaluation of reduced-order aeroelastic simulations for shock-dominated flows

Aeroelastic predictions of a thin panel in supersonic flow with a shock/boundary layer interaction (SBLI) are evaluated relative to wind tunnel experiments. Measured time histories of the panel displacement and velocity show stationary deflections as well as post-flutter oscillations for a range of shock impingement strengths. The fully coupled, computationally efficient aeroelastic modeling framework is formulated with a nonlinear structural reduced-order model and quasi-steady enriched piston theory aerodynamics. The aeroelastic model predicts the observed shape of the stationary responses with correlation coefficients larger than 0.8 for all SBLI conditions. In regards to dynamic responses, the model captures flutter onset as well as the post-flutter periodic oscillations for attached SBLI. The system’s sensitivity to different initial conditions and shock impingement location is also investigated for attached SBLI. Shock-induced flow separation emerges as the SBLI strength increases. For the strongest shock, the computations predict flutter but do not capture the frequency content of the measured post-flutter response. This discrepancy is tied to the breakdown of the quasi-steady flow assumption in piston theory with increasing flow separation and highlights the need for unsteady flow models.

Shock interaction induced heat flux augmentation in hypersonic flows

This paper gives a summary of dedicated experiments on the shock interaction induced heat flux augmentation, by means of tests carried out in the hypersonic wind tunnel H2K. The first test case is devoted to the shock boundary layer interaction on a flat plate. The interaction impact has been varied by changing the free stream parameters and the position of the shock generator, i.e. shock impingement point on the plate. The heat flux distribution has been determined using surface temperature data measured by an infrared camera. The heat flux data combined with free stream flow parameters allow calculation of the Stanton number evolution. The second test case is a double sphere configuration with a variable axial and lateral distance between the spheres. This allowed measurements of the heat flux augmentation induced by a shock-shock interaction along the complete frontal surface of the second sphere, which was hit by the bow shock of the first sphere. Shock-shock and shock-boundary layer interaction effects are studied by means of experiments on the IXV flight configuration with double control flaps. Depending on the test configuration and flow parameters, shock interaction induced heat flux augmentation factors up to seven have been measured.Graphical abstract

Transitional Shock Boundary Layer Interactions on a Compression Ramp at Mach 4

Strong laminar/transitional shock boundary layer interactions (SBLIs) have been investigated in a Mach 4 vacuum-driven wind tunnel with supporting computational fluid dynamics analysis. Such flows are extremely susceptible to large-scale separation, unsteadiness, and high surface heat flux, which can limit control authority on high-speed vehicles. The research community has heavily focused on turbulent interactions, leaving little understanding of how laminar/transitional cases behave with varying external parameters. Four compression ramp angles of 15, 18, 22, and 28° have been tested on a flat plate (unit Reynolds number ) with a range of external length scales (flat plate Reynolds number varies between and at the ramp corner). Laminar separation was observed at boundary-layer thickness Reynolds number between 2700 and 4300, with the centerline length of separation scaling mildly with Reynolds number, increasing proportionally with . Transitional reattachment is observed for strong SBLIs at higher values of . Analysis of high-speed schlieren images shows that the separation shock exhibits low-frequency unsteadiness at a Strouhal number near , consistent with turbulent SBLIs. Detailed phase analysis of schlieren data shows no evidence of upstream influence. Instead, motion of the shear layer and reattachment shock precedes motion of the separation shock foot, suggesting a downstream unsteadiness mechanism. In addition, extensive analysis of subtle features within the interaction identified a slow-moving density disturbance within the bubble that convects toward the shock foot and directly leads separation shock motion. Negligible coherence was observed between low-frequency separation shock motion and acoustic content within the separation bubble.

Prediction of aerothermal characteristics of a generic hypersonic inlet flow

The accurate prediction of aerothermal surface loading is of paramount importance for the design of high speed flight vehicles. In this work, we consider the numerical solution of hypersonic flow over a double-finned geometry, representative of the inlet of an air-breathing flight vehicle, characterized by three-dimensional intersecting shock-wave/turbulent boundary-layer interaction at Mach 8.3. High Reynolds numbers ($Re_L \approx 11.6 \times 10^6$ based on free-stream conditions) and the presence of cold walls ($T_w/T_o \approx 0.27$) leading to large near-wall temperature gradients necessitate the use of wall-modeled large-eddy simulation (WMLES) in order to make calculations computationally tractable. The comparison of the WMLES results with experimental measurements shows good agreement in the time-averaged surface heat flux and wall pressure distributions, and the WMLES predictions show reduced errors with respect to the experimental measurements than prior RANS calculations. The favorable comparisons are obtained using an LES wall model based on equilibrium boundary layer approximations despite the presence of numerous non-equilibrium conditions including three dimensionality, shock-boundary layer interactions, and flow separation. Lastly, it is also demonstrated that the use of semi-local eddy viscosity scaling (in lieu of the commonly used van Driest scaling) in the LES wall model is necessary to accurately predict the surface pressure loading and heat fluxes.

Control of the oscillations of shock train using boundary layer suction

The isolator plays a major role in the scramjet engine by providing suitable pressure for the combustion process. It is situated between the inlet and the combustion chamber. The deceleration of supersonic flow into subsonic flow is achieved by a series of bifurcated normal shock waves referred to as “shock train”. The flow field is more complex with shock–shock interaction and shock boundary layer interaction. The presence of such flow inside the isolator can degrade the performance of the scramjet engine. The present study focuses on the self-excited shock train flow in an isolator and its control by partial removal of the boundary layer. The results examine the variation of the inlet to outlet pressure ratio with and without suction flow control. Numerical results indicate that the boundary layer suction will cause downstream movement of the shock train and the length of the shock train is reduced. The transformation of the shock train into a single curved normal shock is observed when the shock train present closer or upstream of the suction slot region. The obtained result shows the self-excited oscillation frequency in the range of 14-15 Hz. The suction control provides an advantage by 50% reduction in the amplitude of the shock train movement.

Spatially developing supersonic turbulent boundary layer subjected to static surface deformations

The effects of static surface deformations on a spatially developing supersonic boundary layer flow at Mach number M=4 and Reynolds number Reδin≈49300, based on inflow boundary layer thickness (δin), are analyzed by performing large eddy simulations. Two low-order structural modes of a rectangular clamped surface panel of dimensions ≈33δin×48δin are prescribed with modal amplitudes of δin. The effects of these surface deformations are examined on the boundary layer, including changes in the mean properties, thermal and compressibility effects and turbulence structure. The results are analyzed in the context of deviations from concepts typically derived and employed for equilibrium turbulence. The surface deflections, to some degree, modify the correlations that govern both Morkovin’s hypothesis and strong Reynolds analogy away from the wall, whereas in the near-wall region both the hypotheses breakdown. Modifications to the turbulence structure due to the surface deformations are elucidated by means of the wall pressure two-point correlations and anisotropy invariant maps. In addition to the amplification of turbulence, such surface deformations lead to local flow separation, instigating low-frequency unsteadiness. One consequence of significance to practical design is the presence of low frequency unsteadiness similar to that encountered in impinging or ramp shock boundary layer interactions.

Experiments on a Thin Panel Excited by Turbulent Flow and Shock/Boundary-Layer Interactions

An experimental methodology was designed to study the effects of turbulent flow and shock/boundary-layer interaction (SBLI) on the postflutter response of a thin, buckled panel. The approach combines a systematic testing strategy with state-of-the-art full-field, noncontact measurement techniques, including pressure-sensitive paint, three-dimensional digital image correlation (DIC), and schlieren imaging. These measurements, combined with traditional sensors, were used to characterize the loading conditions and the structural response. The thermally buckled panel experienced bistable, cross-well oscillations with cavity pressure modulation for turbulent flow conditions. This response was characterized as chaotic using several conventional criteria. As the temperature differential between the frame and panel decreased, the panel response transitioned to periodic. For the first time, shock-induced limit cycle oscillations (LCOs) were also measured using DIC. In the presence of an attached SBLI, the dynamic response was centered about the three-quarter chord. As the strength of the SBLI increased, the LCO initially subsided and then reappeared near the midchord for the strongest shock considered.

Investigation on shear layer instabilities and generation of vortices during shock wave and boundary layer interaction

The interaction of a reflected shock with the boundary layer inside the shock tubes is an important engineering problem. Studies related to shock wave mitigation and attenuation are performed inside the shock tubes. A proper understanding of the flow behind the reflected shock and the separated zone involving multiple vortical structures is highly essential for estimating the effectiveness of the shock wave mitigation/ attenuation. Such complex flows consist of Lambda shock, shear layer originating from the triple point, multiple shocklets, Mach stems, and vortices. Experimentally the shock structures are obtained through optical techniques. The vortices present in the compressible flow can be obtained through numerical simulations. The complex flows consisting of the above- mentioned features have been simulated numerically so far for the Reynolds numbers up to 1000 [ Zhou G, Xu K, Liu F. Grid-converged solution and analysis of the unsteady viscous flow in a two-dimensional shock tube. Phys Fluids 2018;30(016102):1-21.]. In the present investigation, the shock-wave boundary-layer interaction is simulated for the Reynolds numbers of 1000 and 2500 using a 13th order hybrid scheme to discern the distinct flow features. First, the solver [Kundu A, De S. Navier-Stokes simulation of shock-heavy bubble interaction: Comparison of upwind and WENO schemes. Comput Fluids 2017;157:131-145.] was validated with the benchmark wall density data for a Reynolds number of 1000. Next, the simulations were performed using 50 and 109.5 million cells for the Reynolds number of 2500. The density gradients, vorticity, wall density, and Fourier spectra were used for comparing the flow field for the Reynolds numbers of interest. The Lambda shock, Kelvin-Helmholtz (K-H) vortices in the shear layer, shocklets, the height of lambda shock, and Mach stems were obtained using a grid-mesh of 109.5 million cells. It is observed that the number of vortices generated inside the separated flow region increased with the increase in Reynolds number from 1000 to 2500. Furthermore, the triple point height and the number of K-H vortices generated at the shear layer also increase with an increase in Reynolds number. The present simulations revealed the formation of vortices close to the wall at a Reynolds number of 2500. Such flow structures have an important role in shock / blast wave mitigation and the associated aeroacoustics.

Length scale for the estimation of buzz frequency in the limit of high mechanical blockage in mixed-compression intakes

Oscillatory flow features are common in the unstart of hypersonic mixed-compression intakes and can be classified as low-amplitude or high-amplitude oscillatory unstarted flows. The low-amplitude oscillatory unstarted flow is driven by the shear layer from shock interactions ahead of the cowl, while the high-amplitude oscillatory unstarted flow is driven by the separation caused by shock–boundary-layer interaction on the ramp. While previous studies have observed these flow features and reported their associated frequency, there is no simple criterion available for predicting which mode will occur, and there is a lack of consensus on the appropriate frequency scaling parameter. We study a mixed-compression hypersonic intake in a hypersonic wind tunnel by varying the internal contraction ratio and the throttling ratio to observe various kinds of unstart regimes. Two significant conclusions emerge from considering the results for high-throttling-ratio conditions $(TR > 0.55)$ from the current as well as previous studies. Firstly, the actual shock-on-lip condition at the cowl corresponding to the unthrottled condition, as observed from schlieren images, demarcates the boundary between the two modes of oscillatory unstart flows upon throttling. Secondly, a suitable length scale $(l^*)$ , defined as the extent of the subsonic region in the unstarted flow (as observed from the experimental schlieren images), gives the appropriate frequency scaling parameter ( $f^* = a_0/4l^*$ where $a_0$ is the stagnation acoustic speed).

Application of a long short-term memory neural network for modeling transonic buffet aerodynamics

In the present work, a reduced-order modeling (ROM) framework based on a long short-term memory (LSTM) neural network is applied for the prediction of transonic buffet aerodynamics. This type of network has a high potential for modeling sequential data, which is favorable for capturing the time-delayed effects associated with unsteady aerodynamics. Therefore, the nonlinear identification procedure as well as the generalization of the resulting ROM are presented. Further, a Monte-Carlo-based training procedure is performed in order to estimate statistical errors. The training data set for the ROM is provided by means of forced-motion unsteady Reynolds-averaged Navier Stokes (URANS) simulation. Subsequent to the training process, the ROM is applied for the computation of time-varying integral quantities such as aerodynamic force and moment coefficients. The most challenging aspect when considering buffet aerodynamics is given by the reproduction of the self-sustained unsteadiness of the buffeting flow. Even without any external excitation, the flow is characterized by large shock-boundary layer interaction, resulting in shock movement and flow separation. Finally, the performance of the trained network is demonstrated by predicting the aerodynamic loads of the NACA0012 airfoil considered at transonic freestream conditions. Therefore, the airfoil is excited by a forced pitching motion beyond the buffet-critical angle of attack. A comparison with a full-order computational fluid dynamics (CFD) solution shows that the essential characteristics of the nonlinear buffet phenomenon are captured by the ROM method.

An implicit kinetic inviscid flux for predicting continuum flows in all speed regimes

In this study, the kinetic inviscid flux (KIF) is improved and coupled with an implicit strategy. The KIF is a recently proposed numerical method, which is a dynamic combination of the kinetic flux vector splitting (KFVS) method and the totally thermalized transport (TTT) method. The inherent microscopic mechanism of the KFVS makes the KIF good at solving shock waves and avoiding the numerical shock instability phenomenon. When developing the implicit KIF, it is noticed that, in boundary layers, the KFVS part of the KIF not only reduces the accuracy but also seriously reduces the Courant–Friedrichs–Lewy (CFL) number. As a result, a new weight is proposed in this paper to combine the KFVS method with the TTT method properly. Besides admitting the use of larger CFL numbers, this new weight also contributes to more accurate numerical results like pressure, friction coefficient, and heat flux when solving shock waves, boundary layers, and complex supersonic/hypersonic flows. To examine the validity, accuracy, and efficiency of the proposed method, six numerical test cases covering the whole speed regime are conducted, including the hypersonic viscous flow past a cylinder, the hypersonic double-cone flow, the hypersonic double-ellipsoid flow, the laminar shock-boundary layer interaction, the supersonic flow around a ramp segment and the subsonic lid-driven cavity flow.

Topological modifications due to ramped vanes in a flare-induced shock–boundary layer interaction flowfield

Effect of ′ramped vane′-type vortex generators on a shock-induced flow separation in the vicinity of an axisymmetric compression corner was evaluated. Numerical simulations were performed at Mach 2 on a cone–cylinder–flare model with a flare angle of 24°. The undisturbed boundary layer thickness (δ) at the location of the compression corner was 5 mm. A single array of these vortex generators with a device height of 0.28δ (1.4 mm) was placed on the cylinder surface at different streamwise positions, viz. 5δ, 10δ and 15δ upstream of the compression corner, and their ability to manipulate the shock–boundary layer interaction flowfield was compared. The presence of these devices caused substantial changes in the interaction region and the separation bubble structure. The separation bubble transformed into a series of spade-shaped structures with pockets of attached flow in between them. The ramped vanes increased the separation length along the device centreline, but this effect was attenuated considerably, by bringing them closer to the interaction region. Moving the ramped vanes closer also had a collapsing effect on the spade-shaped structures, which simultaneously widened the attached flow zones in between them.

High-Speed PIV of shock boundary layer interactions in the transonic buffet flow of a compressor cascade

The dynamics of a transonic buffet flow on the suction side of a highly loaded 2-D compressor blade model is investigated at a chord-based Reynolds number of 1.4 times 10^6 and an inlet Mach number of 1.05. Near stall the detached bow shock exhibits pronounced modal shock oscillations at buffet frequencies of {omega ^*}=2pi f c / U_1=2.27 which are not related to any structural aeroelastic modes. High-speed PIV at several stations along the chord provides chordwise velocity spectra and wave propagation velocities of shock-induced perturbations. For this purpose, a double-pulse laser system with a high-repetition rate was set up consisting of two combined DPSS lasers. This enables time-resolved PIV using frame straddling at up to 100 kfps and pulse separations down to 800 ns. Synchronized high-speed shadowgraph imaging simultaneously locates the position of the bow shock. Based on cross-correlations between velocity time records at two points and between velocity and shock position, the propagation velocity of the modal perturbations is determined upstream and downstream from the shock. The measured data indicate that feedback occurs between a region immediately downstream of the shock foot and a plane upstream of the shock, up to which transverse velocity disturbances are convected. This observation is contrary to Lee’s model Lee (AIAA J 28(5):942–944, 1990) which describes self-sustained shock-buffet as a consequence of shock–trailing edge interactions.Graphic

STREAmS: A high-fidelity accelerated solver for direct numerical simulation of compressible turbulent flows

We present STREAmS, an in-house high-fidelity solver for direct numerical simulations (DNS) of canonical compressible wall-bounded flows, namely turbulent plane channel, zero-pressure gradient turbulent boundary layer and supersonic oblique shock-wave/boundary layer interaction. The solver incorporates state-of-the-art numerical algorithms, specifically designed to cope with the challenging problems associated with the solution of high-speed turbulent flows and can be used across a wide range of Mach numbers, extending from the low subsonic up to the hypersonic regime. From the computational viewpoint, STREAmS is oriented to modern HPC platforms thanks to MPI parallelization and the ability to run on multi-GPU architectures. This paper discusses the main implementation strategies, with particular reference to the CUDA paradigm, the management of a single code for traditional and multi-GPU architectures, and the optimization process to take advantage of the latest generation of NVIDIA GPUs. Performance measurements show that single-GPU optimization more than halves the computing time as compared to the baseline version. At the same time, the asynchronous patterns implemented in STREAmS for MPI communications guarantee very good parallel performance especially in the weak scaling spirit, with efficiency exceeding 97% on 1024 GPUs. For overall evaluation of STREAmS with respect to other compressible solvers, comparison with a recent GPU-enabled community solver is presented. It turns out that, although STREAmS is much more limited in terms of flow configurations that can be addressed, the advantage in terms of accuracy, computing time and memory occupation is substantial, which makes it an ideal candidate for large-scale simulations of high-Reynolds number, compressible wall-bounded turbulent flows. The solver is released open source under GPLv3 license. Program summaryProgram Title: STREAmSCPC Library link to program files:https://doi.org/10.17632/hdcgjpzr3y.1Developer’s repository link:https://github.com/matteobernardini/STREAmSCode Ocean capsule:https://codeocean.com/capsule/8931507/tree/v2Licensing provisions: GPLv3Programming language: Fortran 90, CUDA Fortran, MPINature of problem: Solving the three-dimensional compressible Navier–Stokes equations for low and high Mach regimes in a Cartesian domain configured for channel, boundary layer or shock-boundary layer interaction flows.Solution method: The convective terms are discretized using a hybrid energy-conservative shock-capturing scheme in locally conservative form. Shock-capturing capabilities rely on the use of Lax–Friedrichs flux vector splitting and weighted essentially non-oscillatory (WENO) reconstruction. The system is advanced in time using a three-stage, third-order RK scheme. Two-dimensional pencil distributed MPI parallelization is implemented alongside different patterns of GPU (CUDA Fortran) accelerated routines.

Features of panel flutter response to shock boundary layer interactions

The compounding effects of panel flutter and oblique shock impingement are of great concern to the development of light-weight, high-speed vehicles. Shock-induced panel flutter response is investigated at M=2 and Re=120,000 using the Navier–Stokes equations closely coupled to the von-Kármán equations. A naturally occurring laminar inflow boundary layer and incident oblique shockwave are specified through the boundary conditions. The shockwave impinges at the midpoint of a flexible panel. Several incident shock strengths and non-dimensional dynamic pressures are examined. The resulting fluid–structure interaction is significant and is described in detail. In general, the incident shockwaves lead to higher frequency, higher mode flutter when compared to the configuration without an impinging shockwave or with an impinging shockwave in inviscid flow. Additionally, the presence of a compliant surface is shown to promote turbulent transition downstream of the shock boundary layer interaction.

Microvortex Generator Controlled Shock–Boundary Layer Interactions in Hypersonic Intake

AbstractThe occurrence of shock and boundary layer interactions (SBLIs) in aircraft-engine intakes often results in poor total pressure recovery, a degraded boundary layer, and, in worst cases, flo...

Investigation of shock dynamics in an axisymmetric inlet/isolator with attached boundary layers

Abstract