Supersonic Boundary Layer Research Articles

Direct numerical simulation of a supersonic turbulent boundary layer with hydrogen combustion

Direct numerical simulation of a supersonic turbulent boundary layer with hydrogen combustion

Numerical simulation of the interaction between weak shock waves and supersonic boundary layer on a flat plate with the blunt leading edge

The interaction of weak shock waves in form of an N-wave with the supersonic laminar boundary layer on a flat plate with blunt leading edge at the free-stream Mach number M = 2.5 is numerically studied. The numerical results are compared with known experimental data. The combined influence of the N-wave and the leading edge bluntness on the laminar-turbulent transition process is discussed.

Momentum and kinetic energy transport in supersonic particle-laden turbulent boundary layers

Momentum and kinetic energy transport in supersonic particle-laden turbulent boundary layers

The Law of the Total Pressure Loss in Supersonic Streamwise Corner Flow

Abstract The total pressure loss is a critical issue in the design of internal flow channels for hypersonic vehicles. In the flow through the streamwise corner region, the adjacent walls induce secondary flows that exacerbate the viscous effects, thereby intensifying the total pressure loss. This study investigates the impact of the corner’s dihedral angle and the compressibility of the flow on the total pressure loss in the streamwise corner region. The results indicated that as the dihedral angle increases, the distribution of the total pressure contour in the corner region became significantly distorted. The proportion of the boundary layer area within the corner region concurrently increased as the dihedral angle deviates from 90 degrees. Furthermore, we deduce that the total pressure loss is the largest in the supersonic flat plate boundary layer. This paper normalized the total pressure loss coefficient for different dihedral angles in the corner region and established an empirical relationship for the distribution of normalized total pressure loss coefficients along the flow path. It is discovered that under supersonic inflow conditions, the distribution law of the total pressure loss in the corner region exhibits higher predictive accuracy.

Direct numerical simulation of skin friction drag reduction on supersonic turbulent boundary layers with micro-blowing

The flow control mechanism and skin friction drag reduction characteristics of micro-blowing on a Ma2.25 supersonic turbulent boundary layers are investigated through direct numerical simulations, and the effects of blowing intensity and micro-hole arrangement on turbulent structure and skin friction drag in the local control region and downstream region are considered. The results show that the skin friction drag decreases remarkably in the control region under the influence of micro-blowing, and a certain drag reduction can still be maintained in the downstream region. The drag reduction performance in the control region is jointly determined by blowing intensity and micro-hole arrangement. The drag reduction performance of the staggered arrangement is 5.7% and 11.1% higher than that of the inline arrangement at blowing intensities of 0.2% and 0.5%, respectively. However, it is found that the drag reduction in the downstream region is only determined by the blowing intensity and almost independent of the micro-hole arrangement. The effect of micro-blowing on turbulent structures is quantitatively characterized by energy spectrum analysis, and it shows that the streamwise scales of the near-wall streaks are significantly reduced under the influence of micro-blowing. In addition, the compressibility of fluids and the local reverse transfer in the strong expansion region are significantly improved under the influence of micro-blowing. These effects should be considered when performing Large Eddy Simulation modeling of supersonic turbulent boundary layers with micro-blowing.

Slip effects on the receptivity of supersonic flat-plate boundary layer to freestream acoustic waves

The receptivity phase located upstream from the neutral point might be significantly affected by local rarefaction effects (especially surface slip effects) in terms of the boundary-layer transition of near-space hypersonic vehicles. In this paper, the receptivity of a supersonic flat-plate boundary layer to freestream acoustic waves in no-slip and slip flows is analyzed using direct numerical simulations and linear stability theory. The Maxwell–Smoluchowski velocity-slip and temperature-jump boundary conditions are adopted at the wall to account for surface slip effects. A Mach 4.5 flow at different wall-cooling degrees is mainly analyzed, and another Mach-6 case is presented, both with freestream unit Reynolds number on the order of 1×106/m. The main goal is to clarify the qualitative and quantitative influence of surface slip effects on the receptivity phase under different conditions. The results show that the receptivity mechanism in the slip flow is similar to that in the no-slip flow. That is, the mode S or F is excited near the leading edge due to synchronization with slow or fast acoustic waves, and the Mack second mode is excited further downstream after synchronization between modes S and F. However, the slip effects lead to distinctly quantitative differences in receptivity. The slip effects have little influence on the excitation of mode S or F near the leading edge but largely affect the evolution (intermodal exchange) of modes S and F as propagating downstream. Consequently, as for the receptivity to slow acoustic waves, the slip effects play a stabilizing role in receptivity when mode S is stable while a destabilizing role when mode S converts to the first mode in the upstream. As for receptivity to fast acoustic waves, as slip degree increases, the slip effects initially stabilize and then destabilize the receptivity, where the receptivity coefficient of the tested slip case can increase by 25% compared with the no-slip case.

Study on secondary motions in supersonic boundary layers of a bent pipe

The present study employed direct numerical simulation to investigate the supersonic flow of Mach 3 in a bent pipe with a curvature of 0.0825, elucidating the dynamic mechanism of secondary motions within the turbulent boundary layer. The findings indicate that the compressible flow, affected by the wall curvature, is differentiated into several motion patterns as the bending angle increases: a portion of the outer fluid close to the wall, driven by the circumferential pressure gradient, moves inward through the lateral wall, causing an increase in the mass rate toward the lateral boundary layer and promoting the circumferential transport of energy and vorticity; other outer fluids at the start of the bent section, due to the centrifugal force, approach the wall to form a thinner boundary layer downstream; meanwhile, the fluid near the inner wall experiences the expansion, followed by the flow separation and reattachment at a bending angle of 14.6° and 22.0°, respectively, which induce a shear layer that develops from the inner end point toward the mainstream center, gradually reshaping the high-speed flow area within the pipe cross section into a U-shape, and enhancing the vorticity and temperature field of the inner region. Additionally, this study reveals a remarkable phenomenon that the separated flow in a localized inner region forms a rotating field, inducing vortices distinct from the mainstream Dean vortices in the low-speed flow region enclosed by the shear layer.

Linear stability analysis of compressible boundary layer over an insulated wall using compound matrix method: Existence of multiple unstable modes for Mach number beyond 3

The linear spatial stability of a parallel two-dimensional (2D) compressible boundary layer on an adiabatic plate is investigated by considering both 2D and three-dimensional (3D) disturbances. The compound matrix method is employed here, for the first time, for compressible flows, which, unlike other conventional techniques, can efficiently eliminate the stiffness of the equations governing the spectral amplitudes. The method is first validated with published results in the literature corresponding to spatial and temporal instability of flows ranging from low subsonic to high supersonic Mach numbers (M), which shows a good match depending upon the proper choice of free-stream temperature and the wall dispersion relation. Subsequently, flow compressibility effects and the spanwise variation of disturbances are also investigated for M ranging from low subsonic to high supersonic cases (from M = 0.1 to 6). Mack (AGARD Report No. 709, 1984) reported the existence of two unstable modes for M > 3 from viscous calculations (the so-called “second mode”) that subsequently fuse to create only one unstable zone when M increases. Our calculations show a series of higher-order unstable modes for M > 3 in addition to the findings of Ma and Zhong [“Receptivity of a supersonic boundary layer over a flat plate. Part 1. Wave structures and interactions,” J. Fluid Mech. 488, 31–78 (2003)], where such higher-order modes for supersonic boundary layers are all noted to be spatially stable. The number and the frequency extent of the corresponding unstable zones increase with an increase in M beyond 3 while propagating downstream at a higher speed than those corresponding to incompressible, subsonic, and low supersonic (M < 2) cases.

Feasible suggestions of scale-adaptive simulation approach based on the high-fidelity large-eddy simulation of supersonic turbulent boundary layer

Feasible suggestions of scale-adaptive simulation approach based on the high-fidelity large-eddy simulation of supersonic turbulent boundary layer

Supersonic turbulent boundary layer on a plate. III. Laws of the wall for velocity and temperature

Supersonic turbulent boundary layer on a plate. III. Laws of the wall for velocity and temperature

Generation of inflow turbulence using an improved synthetic eddy method

The generation of turbulent inflow conditions is a key issue in large eddy simulation (LES) or direct numerical simulation (DNS). In this paper, an improved synthetic eddy method (SEM) is proposed to generate inflow turbulence for LES and DNS. The improvements about SEM focus on the eddy radius and eddy distributions. First, the eddy radius is improved to reduce the nonphysical vortex structure on the wall caused by overestimation of the radius. Second, a sampling method using Gaussian distribution is proposed to improve the distribution of eddies, which accurately captures the randomness of turbulent structure size and is close to the actual flow field. The improved method is applied to the direct numerical simulation of the supersonic turbulent boundary layer at Mach 2.7 and the 24° compression ramp. Results indicate that the predictions yielded by the improved method are in good agreement with both DNS and experimental data. Compared to the original method, the improved method exhibits a more rapid recovery of the friction coefficient and effectively shortens the development distances. The improved SEM has enhanced the efficiency and accuracy of generating inlet turbulence, which can provide inlet turbulence boundary conditions for LES and DNS.

Features of the cross-correlation characteristics of disturbances in the supersonic boundary layer of the plate at the influence of a pair of weak shock waves on the leading edge

Features of the cross-correlation characteristics of disturbances in the supersonic boundary layer of the plate at the influence of a pair of weak shock waves on the leading edge

Direct numerical simulation of supersonic boundary layer transition induced by gap-type roughness

The transition of the supersonic boundary layer induced by roughness is a highly intricate process. Gaining a profound understanding of the transition phenomena and mechanisms is crucial for accurate prediction and control. In this study, to delve into the flow mechanisms of a transition in a supersonic boundary layer induced by the medium gap-type roughness, direct numerical simulation is employed to capture and analyze the transition process. Research indicates that as the flow over the flat plate passes the gap, the spanwise convergence effect leads to the formation of both upper and lower counter-rotating vortex pairs. As the flow progresses, these counter-rotating vortex pairs in the central region exhibit attenuation, with streamwise vortices developing on both sides. At a certain downstream distance, the boundary layer becomes unstable, triggering the formation of streamwise vortex legs. These streamwise vortex legs undergo further evolution, transforming into hairpin vortices and leg-buffer vortices. The formation of the central low-speed zone downstream of the roughness element is mainly attributed to the lift-up effect of the low-speed flow propelled by the central counter-rotating vortex pairs. The low-speed streaks on both sides are primarily influenced by the streamwise vortices. Through a meticulous analysis of the turbulent kinetic energy distribution and its generation mechanisms during the transition phase, this study infers that the primary sources of turbulent kinetic energy are the hairpin vortices, leg-buffer vortices, and their consequent secondary vortices. Combined with modal analysis, the study further elucidates the generation and breakdown of hairpin and leg-buffer vortices.

Reduced-Order Modeling of Steady and Unsteady Flows with Deep Neural Networks

Large-eddy and direct numerical simulations generate vast data sets that are challenging to interpret, even for simple geometries at low Reynolds numbers. This has increased the importance of automatic methods for extracting significant features to understand physical phenomena. Traditional techniques like the proper orthogonal decomposition (POD) have been widely used for this purpose. However, recent advancements in computational power have allowed for the development of data-driven modal reduction approaches. This paper discusses four applications of deep neural networks for aerodynamic applications, including a convolutional neural network autoencoder, to analyze unsteady flow fields around a circular cylinder at Re = 100 and a supersonic boundary layer with Tollmien–Schlichting waves. The autoencoder results are comparable to those obtained with POD and spectral POD. Additionally, it is demonstrated that the autoencoder can compress steady hypersonic boundary-layer profiles into a low-dimensional vector space that is spanned by the pressure gradient and wall-temperature ratio. This paper also proposes a convolutional neural network model to estimate velocity and temperature profiles across different hypersonic flow conditions.

Direct numerical simulation of mechanism and control of secondary instability induced transition in a supersonic boundary layer

Mechanisms and control of secondary-instability-induced-transition in a supersonic boundary layer are studied numerically via direct numerical simulation. The aim is to investigate and compare the transition mechanisms of fundamental, subharmonic, asymmetric subharmonic, and detuned resonances, and to control these secondary instabilities using a local wall cooling strip. The results indicate that the nonlinear interaction between the high-amplitude primary mode and low-amplitude secondary modes is the main contributor to transition. The mutual- and self-interactions of the primary and secondary modes generate other harmonic modes with laminar breakdown soon appearing. The asymmetric subharmonic resonance induces the earliest transition, while the fundamental subharmonic has the latest. Wall cooling effects are also studied. The results show that a lower wall temperature significantly suppresses the secondary instabilities, and steady modes become dominant and lead to obvious streamwise vortexes. Numerical data demonstrate that all secondary-instability-induced transitions result in fully developed turbulent boundary layers, as supported by the skin friction and scaled velocity profiles. The transition control cases indicate that the local wall cooling strip can significantly delay the transition by suppressing the growth of the primary mode. An upstream control strip is found to have a more obvious suppression effect. The fundamental and asymmetric subharmonic resonances are sensitive to the location of the local wall cooling strip and show a stronger transition delaying effect.

Excitation and evolution of radiating modes in supersonic boundary layers. Part 1. Fundamental resonance with impinging sound waves

Excitation and evolution of radiating modes in supersonic boundary layers. Part 1. Fundamental resonance with impinging sound waves

Excitation and evolution of radiating modes in supersonic boundary layers. Part 2. Back effect of spontaneously radiated Mach waves

Excitation and evolution of radiating modes in supersonic boundary layers. Part 2. Back effect of spontaneously radiated Mach waves

Flow structure and cooling effect of seeping gas film in supersonic turbulent boundary layer

Seeping gas film cooling is an active cooling method that involves injecting a cooling gas film into the boundary layer through porous media. In this study, a sintered wire mesh porous material is used to generate the gas film through seepage flow. The flow of the gas film in the supersonic turbulent boundary layer is observed using nanometer particle plane laser scattering technology, while the temperature distribution of the porous wall surface is measured by infrared thermometry. It is observed that as the injection rate of the gas film exceeds 0.2 %, the supersonic turbulent boundary layer thickens significantly. The rate of thickness growth is promoted from 5.5 % to 17.1 % as the injection rate increases from 0.3 % to 0.8 %. Additionally, the injection of the gas film induces an oblique shock wave, and the shock angle increases with the enhancement of the injection rate, resulting in an uneven cooling effect along the injector length. The transient cooling efficiency of the seeping gas film under thermal non-equilibrium conditions is analyzed. The results indicate that the temperature amplitude on the porous wall surface decreases significantly when the injection rate is promoted from 0.2 % to 0.3 %. However, continuing to increase the coolant injection rate does not lead to a more significant improvement in cooling efficiency. When the injection rate reaches 0.5 %, the relative cooling efficiency reaches 50 %–60 %, which is much lower than the predicted values of the thin film theory.

Roughness-Induced Transition in Supersonic Boundary Layers with Wall Heating and Cooling Effects

The laminar–turbulent transition of a supersonic boundary layer induced by an isolated roughness element is investigated using direct numerical simulation, BiGlobal linear stability theory, and the three-dimensional parabolized stability equation. Cylindrical and diamond-shaped roughness elements are investigated in combination with different wall temperatures. Direct numerical simulations show that the cylinder configuration induces an earlier transition than the diamond configuration, with the interaction between the separated shear layer and the counter-rotating vortices causing the transition. BiGlobal analysis and the parabolized stability equation confirm the existence of two unstable instability modes in the wake region, namely, a symmetric mode and an antisymmetric mode, with the former being strongly dominant. The wall cooling and heating effects are studied by changing the wall temperature. Wall heating lifts the inlet boundary layer and weakens the separated shear layer. This, in turn, weakens the wake instability and delays the transition. The antisymmetric model disappears as the wall heating increases. Wall cooling accelerates the transition by enhancing the distortion of the roughness wake, resulting in stronger instabilities. Finally, the Reynolds number based on the momentum defect is used to define the transition criterion, and this is found to be in good agreement with the present simulation results.

Numerical Simulation of the Interaction between Weak Shock Waves and Supersonic Boundary Layer on a Flat Plate with the Blunt Leading Edge

The interaction of weak shock waves in form of an N-wave with the supersonic laminar boundary layer on a flat plate with blunt leading edge at the free-stream Mach number M = 2.5 is numerically studied. The numerical results are compared with known experimental data. The combined influence of the N-wave and the leading edge bluntness on the laminar-turbulent transition process is discussed.