Supersonic Boundary Layer Transition Research Articles

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.

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.

Effect of a streamwise vortex on supersonic boundary layer transition

Supersonic boundary layer transition induced by a streamwise vortex is experimentally investigated based on the nanoparticle-based planar laser scattering technology, with an inflow Mach number of 2.7 and a unit Reynolds number of 8.79 × 106 m−1. Though the vortex is outside the boundary layer, it results in obvious difference of transition position along spanwise direction. On the up-wash side, transition to turbulence is completed in advance compared to the vortex-free case, and transition position moves more upstream as the vortex approaches the wall. Conversely, the transition process on down-wash side is inhibited. The promotion and inhibition effects on transition are found to be induced by the influence of vortex on streak density.

Time-resolved visualization of coherent structures during supersonic boundary layer transition

The coherent structures arising during flat-plate boundary layer transition at Mach number 3.4 are investigated using a custom-built hyper-rate imaging system. The evolution of transitional structures is investigated in the Eulerian and Lagrangian reference frames. The upstream evolution of transition is dominated by the generation of new hairpin structures, while the interaction among multiple structural types dominates the evolution downstream. The breakdown of the existing structure, which may be caused by interactions among multiple types of structures with similar scales, is also visualized.

Experimental Study of Supersonic Boundary Layer Transition Induced by Cylindrical Roughness Elements

Experimental Study of Supersonic Boundary Layer Transition Induced by Cylindrical Roughness Elements

Numerical study of the suction flow control of the supersonic boundary layer transition in a framework of gas-kinetic scheme

In this paper, we numerically study the supersonic boundary layer transition process and the suction flow control method by employing the DNS technology. By adopting the high resolution GKS method, the mechanism of the boundary transition and the statistical characteristics of the compressible turbulent boundary layer are investigated. Results indicate that the streamwise vortex and the three-dimensional shear layer with high instability are the main causes of the boundary layer transition. Suction control cases show that the laminar boundary layer suction can weaken the development of the unstable disturbance waves in the boundary layer, and delay the streamwise position of the local disturbance. Also, the space distributions of the vortex structure are weakened to delay the process of the transition.

Influence of the Entropy Layer on Turbulence Transition of the Supersonic Boundary Layer in the Wake with Isolated Roughnesses

The influence of the entropy layer generated by a small bluntness of the flat plate leading edge on laminar-turbulent supersonic boundary layer transition in the wake behind isolated roughness elements located on the plate surface is studied when the bluntness radius is varied. The experiments are carried out at the Mach number M = 6 over the Reynolds number range (0.4–2.3) × 106. The Reynolds number is calculated on the basis of the free-stream parameters at the point of roughness. The well-known correlation dependences are verified and the transition criterion is estimated.

Numerical study on supersonic boundary-layer transition and wall skin friction reduction induced by fuel wall-jet combustion

In this study, the effect of wall-jet combustion on boundary layer transition and skin friction reduction was numerically investigated. To effectively capture the characteristics of boundary layer flow, the transition k−kl−w model was employed as the turbulence model and laminar finite-rate model was chosen as the combustion model. This numerical method was firstly validated by two sets of experimental results in open domain. After that, the research on wall-jet combustion was conducted and the numerical results showed that when injecting hydrogen in different directions, both the hydrogen self-ignition location and the skin friction on the wall were not altered significantly. When the injection angle to the airflow direction was increased to 30°, the intensity of combustion was insufficient and the skin-friction coefficient would be increased. Meanwhile, boundary layer transition occurred in a relatively smaller Reynolds number at this condition. The variation of the wall-jet height would have a greater impact on both boundary layer transition and the skin friction. More components can diffuse to the lower wall as the height of the wall-jet was enlarged, which can make the process of boundary layer transition be postponed and the skin friction be reduced as well. Furthermore, greater skin-friction reduction would be achieved downstream the self-ignition location when the wall was adiabatic, while the original Reynolds numbers for boundary layer transition is the smallest if the wall temperature is set to 600 K. Finally, in order to simulate the effect of the back pressure on the combustion flow field, another injector is added near the exit at the upper wall to produce air-throttling flow jet. The results showed that altering back pressure nearly has little influence on boundary layer transition and the skin friction decreases with more air-throttling flow rate.

Experimental study on supersonic plate boundary layer transition under pulsed arc plasma excitation control

Pulsed arc plasma excitation is characterized by strong local heating effect and wide disturbance range, and it has a broad application prospect in supersonic flow control. In this paper, by using electrical parameter measurement system and high speed schlieren technique, we study the electrical and flow field characteristics of pulsed arc plasma excitation under the condition of <i>Ma</i> = 3 incoming flow. The nano-particle planar laser scattering (NPLS) is used to investigate the flow structure of the supersonic flat boundary layer, and the transition characteristics of the boundary layer at different plasma excitation frequencies are studied. The experimental results show that in the flow field with <i>Ma</i> = 3 and the total incoming pressure <i>P</i><sub>0</sub> = 1 atm (1 atm = 1.01 × 10<sup>5</sup> Pa), the peak voltage of the pulsed arc plasma actuator discharge is 6 kV, the peak current is 70 A, the time scale of the discharge is about 300 ns, the single discharge energy is 70 mJ; the pulsed arc discharge will produce the precursor shock wave with higher velocity and the thermal deposition zone with higher temperature, which will exert continuously disturbance on the boundary layer. The pulsed arc plasma excitation with perturbation can promote the transition of supersonic plate boundary layer. Moreover, the high-frequency impact effect of pulsed discharge can promote the transition to occur ahead of time, and the higher the frequency, the better the effect is. As the excitation frequency increases, the transition position of the boundary layer of the supersonic flat plate moves forward, and the length of the transition area of the boundary layer becomes shorter as the excitation frequency increases. When the excitation frequency is 60 kHz, the length of transition zone is 0 and the thickness of turbulent boundary layer is 25 mm. When a high frequency is applied (<i>f</i> = 40, 60 kHz), the transition path of the boundary layer is that the shock wave generated by the plasma excitation triggers the unstable wave, and the development of unstable waves directly skips the linear growth stage, passes through the bypass and transitions into turbulent flow. The pulsed arc plasma excitation can be used to promote supersonic boundary layer transition.

Supersonic boundary layer transition induced by self-sustaining dual jets**Project supported by the National Natural Science Foundation of China (Grant Nos. 11602299, 11872374, and 51809271).

To promote high-speed boundary layer transition, this paper proposes an active self-sustaining dual jets (SDJ) actuator utilizing the energy of supersonic mainflow. Employing the nanoparticle-based planar laser scattering (NPLS), supersonic flat-plate boundary layer transition induced by SDJ is experimentally investigated in an Ma-2.95 low-turbulence wind tunnel. Streamwise and spanwise NPLS images are obtained to analyze fine flow structures of the whole transition process. The results reveal the transition control mechanisms that on the one hand, the jet-induced shear layer produces unstable Kelvin–Helmholtz instabilities in the wake flow, on the other hand, the jets also generates an adverse pressure gradient in the boundary layer and induce unstable streak structures, which gradually break down into turbulence downstream. The paper provides a new method for transition control of high-speed boundary layer, and have prospect both in theory and engineering application.

Experimental study of the laminar-turbulent transition on models of wings with subsonic and supersonic leading edge at M = 2

The paper is devoted to an experimental study of laminar-turbulent transition in supersonic boundary layers on swept wings at Mach 2. The experiments were fulfilled in the low nose supersonic wind tunnel T-325 of Khristianovich Institute of Theoretical and Applied Mechanics on the models with supersonic or subsonic leading edges. The transition location was determined using a constant temperature hot-wire anemometer. while experiments for the case of a supersonic leading edge are a continuation of previous studies, then for the case of a subsonic leading edge, the first data have been obtained that show the possibility of carrying out studies on the development of unstable perturbations. It was obtained that transition take place parallel to the leading edge for both models. The destabilizing effect of the surface curvature of the model on the position of the laminar-turbulent transition for the case of a supersonic leading edge was confirmed. This experimental data is the room for further investigation.

Experimental investigation of the natural and controlled disturbance development in a supersonic boundary layer on the swept wing

The results of investigation of the transverse flow inhomogeneity effect on the natural and controlled disturbance development in 3D supersonic boundary layer are considered. The transition curves were measured on a model of a smooth swept wing and on a swept wing with roughness elements of two configurations. The influence of periodic surface roughness on the growth of natural pulsations (without using a discharge) in the boundary layer is determined. In experiments with artificial disturbances it is found that the periodic modulation of mean flow can lead to stabilization of controlled disturbances in supersonic boundary layer on a swept wing. It was found out that the weak flow inhomogeneity can influence the effectiveness of the active/passive technology of supersonic boundary layer transition control, including the surface microroughness on a swept wing.

Data-driven construction of a reduced-order model for supersonic boundary layer transition

In this study, a data-driven method for the construction of a reduced-order model (ROM) for complex flows is proposed. The method uses the proper orthogonal decomposition (POD) modes as the orthogonal basis and the dynamic mode decomposition method to obtain linear equations for the temporal evolution coefficients of the modes. This method eliminates the need for the governing equations of the flows involved, and therefore saves the effort of deriving the projected equations and proving their consistency, convergence and stability, as required by the conventional Galerkin projection method, which has been successfully applied to incompressible flows but is hard to extend to compressible flows. Using a sparsity-promoting algorithm, the dimensionality of the ROM is further reduced to a minimum. The ROMs of the natural and bypass transitions of supersonic boundary layers at $Ma=2.25$ are constructed by the proposed data-driven method. The temporal evolution of the POD modes shows good agreement with that obtained by direct numerical simulations in both cases.

Laminar-turbulent transition of supersonic boundary layer in the presence of external disturbances

A method for estimating the location of the laminar-turbulent transition is described, which is based on the theory of linear stability under the condition of the acoustic field presence. In the stable region near the stability loss location, the amplitude of the velocity disturbances is determined, which are excited by an external sound wave. By virtue of the proximity of sound wave parameters and proper oscillations, the obtained value of the amplitude is accepted to be equal to the amplitude of the growing wave of the boundary layer. At some downstream point, the Reynolds stresses corresponding to this amplitude become equal to several percents of laminar stresses of the undisturbed boundary layer. Based on experimental data it is accepted in the work that at the transition onset, the Reynolds stresses amount to 12 % of laminar stresses.

Effect of thermo-mechanical non-equilibrium on the onset of transition in supersonic boundary layers

Direct numerical simulations (DNS) for supersonic boundary layers (SBLs) with a free-stream Mach number of M∞ = 2.2 are carried out. Various cases are investigated, involving the adiabatic and the isothermal (cooled and heated) walls. The laminar boundary layer is tripped using a blowing and suction strip with single-frequency and multiple spanwise wave-number excitation. Effects of thermo-mechanical non-equilibrium of thermal boundary layer on laminar-to-turbulent transition (LTT) are presented in detail. Cases with two perturbation intensities are investigated (0.5% and 2.4%). The receptivity analysis of transition onset location towards the thermo-mechanical non-equilibrium is performed using different physical quantities like streamwise evolution of skin-friction coefficient, Stanton number and Dynamic mode decomposition (DMD). The results reveal that thermo-mechanical non-equilibrium tends to advance the transition onset location and also decreases the transition length for the heated walls regardless of the initial perturbation intensity. However, for the cooled walls with 2.4% perturbation intensity, the existence of thermo-mechanical non-equilibrium has a stabilizing effect resulting in delayed transition onset. The flow stays laminar for cooled walls with 0.5% perturbation intensity. The results obtained from DMD analysis uncover two distinct ways of evolution for odd and even harmonics of the perturbation frequency. DMD results also show that the fundamental evolution of the modes is not affected by the physical flow parameters like wall temperature or existence of thermo-mechanical non-equilibrium. It is observed that the imposed frequency mode or the principal mode is dominant in the transition region and eventually breakdown to smaller structures in the turbulent regime.

Laminar-to-turbulent transition in supersonic boundary layer: Effects of initial perturbation and wall heat transfer

Direct numerical simulations of supersonic boundary layers (SBLs) over a flat plate for M∞=2.2 are performed for adiabatic and isothermal (cooled and heated) walls. Receptivity analysis based on five criteria, namely skin-friction coefficient, Stanton number, Reynolds shear stress, wall-normal Reynolds heat flux, and modal decomposition are performed. Effect of perturbation intensity and wall heat transfer on the receptivity and the transitional growth of SBLs are presented. It is found out that increasing perturbation intensity moves the transition onset location upstream and increases the transition length. Additionally, below 1% perturbation intensity, wall cooling stabilizes the flow while beyond this value it has the opposite effect.

Laminar-turbulent transition in supersonic boundary layers with surface heat transfer: A numerical study

ABSTRACTThrough high-resolution direct numerical simulations, the present study aims to investigate several laminar-to-turbulent transition scenarios in the presence of wall heat transfer for supersonic boundary layers over strongly heated/cooled and adiabatic flat plates. The laminar boundary layer is tripped using a suction and blowing technique with a single-frequency, multiple-spanwise wavenumber excitation. The results are evaluated and compared with linear stability theory to isolate the effect of wall heat transfer, as well as forcing parameters, on the transition. It was found that increasing the disturbance amplitude as well as perturbation frequency moves the transition upstream. Also, the effect of wall heating was seen to stabilize the flow and to postpone the transition, contrary to the wall cooling.

Numerical Study of Supersonic Boundary-Layer Transition due to Sonic Wall Injection

The boundary-layer transition on a generic hypersonic forebody at Mach 6 downstream of a sonic wall injection is investigated by means of implicit large-eddy simulation, Fourier analysis, and dynamic mode decomposition of numerical data. An academic Mach 4.2 flat-plate configuration with boundary-layer edge conditions matching those of the forebody is considered. Two empirical correlations are proposed to predict the penetration of the underexpanded tripping jet into the boundary layer (that is, the Mach disk height) as a function of the pressure ratio. Different transition mechanisms are observed downstream of the injection port, depending on the jet penetration. The dynamic mode decomposition reveals the spatial structures, temporal frequencies, and growth rates of three-dimensional instability modes. These modes, arising from the jet counter-rotating vortices, are either varicose or sinusoidal, with the latter being more efficient than the former for tripping the boundary layer at low-injection pressure. Recent experiments in the Boeing/U.S. Air Force Office of Scientific Research Mach 6 quiet tunnel at Purdue University show similar transition patterns, but also some discrepancies. A more representative configuration including crossflow effects, and more resolved simulations, is needed for a quantitative comparison.

1stモードが支配的な超音速自然層流翼の 境界層遷移に関する非線形解析

1stモードが支配的な超音速自然層流翼の 境界層遷移に関する非線形解析

A combined experimental and numerical investigation of roughness induced supersonic boundary layer transition

The effect of surface roughness on boundary layer transition is of great importance to hypersonic vehicles. In this paper, both experimental and numerical methods are used to investigate the laminar-turbulent transition of a Mach 3 flat-plate boundary layer induced by isolated roughness element. Good agreements are achieved between experimental data and high-order numerical simulations. It is observed that, with increasing height of roughness, the transition tends to move forward. Two different types of transition mechanisms are found according to the height of the roughness elements. For the smallest roughness height of h=1mm, the shear layer instability in the wake region appears to be the leading mechanism for transition to turbulence. For two larger roughness elements of h=2mm and h=4mm, strong unsteadiness is developed from the upstream separation zone and transition is immediately accomplished, which indicates that the absolute instability in upstream separation zone dominates the transition.