1,720,981 research outputs found
Formation of surface trailing counter-rotating vortex pairs downstream of a sonic jet in a supersonic cross-flow
No full textDirect numerical simulations were conducted to uncover physical aspects of a transverse sonic jet injected into a supersonic cross-flow at a Mach number of 2.7. Simulations were carried out for two different jet-to-cross-flow momentum flux ratios of 2.3 and 5.5. It is identified that collision shock waves behind the jet induce a herringbone separation bubble in the near-wall jet wake and a reattachment valley is formed and embayed by the herringbone recirculation zone. The recirculating flow in the jet leeward separation bubble forms a primary trailing counter-rotating vortex pair (TCVP) close to the wall surface. Analysis on streamlines passing the separation region shows that the wing of the herringbone separation bubble serves as a micro-ramp vortex generator and streamlines acquire angular momentum downstream to form a secondary surface TCVP in the reattachment valley. Herringbone separation wings disappear in the far field due to the cross-interaction of lateral supersonic flow and the expansion flow in the reattachment valley, which also leads to the vanishing of the secondary TCVP. A three-dimensional schematic of surface trailing wakes is presented and explains the formation mechanisms of the surface TCVPs.</p
Turbulence structures and statistics of a supersonic turbulent boundary layer subjected to concave surface curvature
No full textSupersonic turbulent flows at Mach 2.7 over concave surfaces for two different radii of curvature were investigated and compared with a flat plate turbulent boundary layer using direct numerical simulations. The streamwise velocity reduces in the outer part of the boundary layer due to compression, while it increases near the wall due to curvature, with a higher shape factor for the concave cases. The near-wall spanwise streak spacing reduces compared to the flat plate, with large-scale streaks and turbulence amplification also observed. Streamwise velocity iso-surfaces and streamlines show the generation of Görtler-like vortices, consistent with significant centrifugal effects. Abundant small vortices are shown to be associated with large baroclinic production of vorticity that is caused by the density and pressure gradients that are associated with concave compression. Profiles of turbulent kinetic energy and turbulent Mach number exhibit a characteristic two-layer structure in the concave boundary layer cases. In the outer layer, turbulence is greatly amplified, whereas a local balance exists in the inner layer. Turbulent energy budget analysis shows that both production and dissipation increase near the concave wall, whereas in the outer part of the boundary layer, the production is increased and ultimately balanced by convection and turbulent transport.</p
Turbulence decay in a supersonic boundary layer subjected to a transverse sonic jet
No full textThe turbulence state in a supersonic boundary layer subjected to a transverse sonic jet is studied by conducting direct numerical simulations. Turbulence statistics for two jet-to-cross-flow momentum flux ratios of 2.3 and 5.5 based on the previous simulation (Sun & Hu, J. Fluid Mech., vol. 850, 2018, pp. 551-583) are given and compared with a flat-plate boundary layer without a jet. The instantaneous and time-averaged flow features around the transverse jet in the supersonic boundary layer are analysed. It is found that, in the near-wall region, turbulence is suppressed significantly with increasing in the lateral boundary layer around the jet and the turbulence decay is retained in the downstream recovery region. The local boundary-layer thickness decreases noticeably in the lateral downstream of the jet. Analysis of the cross-flow streamlines reveals a double-expansion character in the vicinity of the jet, which involves the reattachment expansion related to the flow over the jet windward separation bubble and the jet lateral expansion related to the flow around the jet barrel shock. The double expansion leads to the turbulence decay in the jet lateral boundary layer and causes a slow recovery of the outer layer in the far-field boundary layer. A preliminary experiment based on the nanoparticle laser scattering technique is conducted and confirms the existence of the turbulence decay phenomenon.</p
Simulation of liquid jet primary breakup in a supersonic crossflow under adaptive mesh refinement framework
No full textCompressible two-phase flows were simulated based on the five-equation model under the Adaptive Mesh Refinement (AMR) framework to balance the requirements between space resolution and computational cost. And the simulation system was established in an open source software AMROC (Adaptive Mesh Refinement Object-oriented C++). A combination of Godunov method and wave propagation method was introduced to integrate numerical methods with the AMR algorithm. High speed and high liquid-gas density ratio are two main challenges in the simulation of liquid jet in a supersonic crossflow. To enhance the robustness of the simulation system, a MOON-type positivity preserving method was adopted in the development of the codes. Based on the system mentioned above, a liquid jet in a Mach 1.5 supersonic crossflow was simulated as the standard case to study the primary breakup process in the near field. The simulation captured the column and surface breakup which were the results of the development of the unstable waves in two directions respectively. The instabilities causing the surface breakup were found to be generated in the transonic region initially. Crossflow of a higher Mach number (Ma 1.8) was found being able to augment the in stable waves along the injection direction and increase the number of instabilities responsible for the surface breakup. While there was no obvious enhancement of the penetration in the condition of periodic injection, extra unstable waves were imposed on both of windward and leeward liquid surface. The introduced unstable waves had an improvement on the column and surface breakup
Direct numerical simulation of a supersonic turbulent boundary layer subject to adverse pressure gradient induced by external successive compression waves
No full textA freestream Mach 2.9 flat-plate supersonic turbulent boundary layer subject to a "pure" adverse pressure gradient (APG) without the impact of wall curvatures is studied by direct numerical simulation and compared with a benchmark flow with zero pressure gradient. Due to APG, the streamwise velocity shows an increase in the near-wall region and a reduction in the outer boundary layer. The principal strain rate shows a sandwich distribution along the wall-normal direction. The mismatch between the temperature and velocity fluctuations in both the inner and the outer layer is observed. Enhanced LSMs (large-scale motions) and large velocity patches are the typical flow structures in the outer and inner boundary layer subject to APG, respectively. From the analysis of quadrant decomposition, the sweep events dominate in the near-wall region while ejection events dominate the rest of the boundary layer. It is found that the baroclinicity plays a significant role in the formation of the enhanced LSMs in the outer boundary layer and the near-wall velocity patches. The resulting amplified vorticity further drives the interactive motions of the outer fluid and inner fluid. The turbulent kinetic energy and turbulent Mach number profiles are amplified by APG and a second peak is observed in both profiles. Turbulent energy budget analysis demonstrates that both the production and viscous effects are strengthened in the near-wall region while in the outer layer, the production is significantly amplified and balanced by the increased convection and turbulent transport.</p
Dynamic detonation stabilization in supersonic expanding channels
No full textIn the present work, dynamic detonation stabilization in expanding channels is numerically investigated by injecting a hot jet into a hydrogen–oxygen combustible mixture flowing at supersonic speed. The two-dimensional reactive Navier–Stokes equations and one-step two-species reaction model are solved using a hybrid sixth-order WENO-Centered Difference (CD) scheme based on the SAMR (Structured Adaptive Mesh Refinement) framework. The results show that the highly unstable shear layer interactions with the unburned jet resulting from the Prandtl-Meyer expansion fan result in numerous large-scale vortices, which contribute significantly to rapid turbulent mixing and diffusion effects. This can further facilitate the consumption of the unburned jet and its subsequent heat release to support the dynamically stationary propagation of detonation. Meanwhile, detonation attenuation in the supersonic flow can be also effectively suppressed because of the formation of a hydrodynamic channel associated with a corresponding hydrodynamic throat. It is indicated that the shear layer interactions with the unburned jet and the generation of hydrodynamic channel can both play important roles in dynamically stationary propagation of detonation in supersonic expanding channels after the shutdown of the hot jet. With the increase of the expansion angle, the enlarged unburned jet is gradually extended out of the sonic line, and the deficit of heat release cannot contribute to stationary propagation of detonation, thus eventually leading to detonation failure. It is indicated that there might exist a critical angle θCT. Dynamic stabilization of detonation can be realized in expanding channels when the angle is smaller than θCT, while the detonation propagates below the CJ velocity and finally fails when the angle is larger than θCT. Through the control of a moving boundary by dynamically changing the expansion angle, the continuous detonation attenuation can be effectively suppressed and finally turned to forward propagation successfully, indicating that dynamically stationary propagation of detonation can be realized through the dynamic control of a moving boundary
Effect of combustion mode on thrust performance in a symmetrical tandem-cavity scramjet combustor
No full textTo investigate the relationship between the thrust performance and combustion mode in a scramjet combustor, the acceleration process of a symmetrical tandem-cavity combustor is numerically investigated based on previous experiments. The numerical results are essentially in agreement with the experimental results through high-speed photography and wall pressure comparisons, which verifies the reliability of the numerical method. Then, three groups of simulations with different equivalent ratios are designed under the condition of inlet flow Mach numbers 2.4, 2.5, 2.6, 2.7, and 2.9. Complete combustion efficiency and specific section thrust are defined to measure the combustion efficiency of kerosene and thrust performance. Four combustion modes are defined and observed in simulations, i.e. strong ram mode, weak ram mode, dual-mode ram mode, and scram mode. The simulation results show that the combustion modes in the former and the later cavities for this tandem-cavity scramjet combustor are related to each other. Combustion modes have a significant impact on the thrust performance of the combustor and the specific section thrust mutation is caused by mode transition. Under different inlet Mach numbers, appropriate equivalent ratio settings can improve thrust performance for this tandem -cavity supersonic combustor. Two ideal combustion modes realized by different injection schemes are proposed for better thrust performance of symmetrical tandem-cavity supersonic combustor.(c) 2022 Elsevier Masson SAS. All rights reserved
Effects of oblique shock waves on turbulent structures and statistics of supersonic mixing layers
No full textA supersonic mixing layer at a convective Mach number of 0.8 was investigated by large eddy simulation. Turbulent structures and statistics of the mixing layer interacting with an oblique shock at different strengths in the self-preserving stage were investigated and compared with the shock-free mixing layer. An inflection point arises on the velocity profiles in the self-preserving region where the incident shock wave impinges, in addition to the three inflection points existing in the shock-free mixing layer. It is caused by the hairpin vortices induced through the baroclinic mechanism of the interaction of the incident shock wave. However, the induced hairpin vortices disappear quickly within a short distance. The vorticity thickness of the shocked-mixing layer experiences a sudden decrease in the vicinity of the shock impingement point, which is due to the induced hairpin vortices, followed by a more rapid growth than that of the shock-free mixing layer. So the incident shock has positive effects on the growth of the mixing layer. Both the hairpin vortices and the vortices originated from the hairpin vortices can result in a double-peak profile of the streamwise Reynolds stress in the transient stage of the mixing layer. In addition, the asymmetric profiles for the Reynolds stress are due to the hairpin vortices breakup earlier in the upper stream. The amplitudes of the Reynolds stress increase slightly and their peak positions move toward the center of the mixing layer even in the self-preserving stage. Moreover, the profiles of the transverse Reynolds stress and Reynolds shear stress have two peaks for the shocked-mixing layer which are caused by the reflected shock waves and the mixing layer. The incident shock increases energy transport and convection between the mixing layer and the mainstream. As a result, the mixing process of the shocked-mixing layer is enhanced.</p
Mechanism of detonation stabilization in a supersonic model combustor
No full textThe present work studies numerically the quasi-steady propagation of a hydrogen/oxygen detonation in a supersonic model combustor consisting of a cavity and an expanding wall. The two-dimensional reactive compressible Navier-Stokes equations with one-step and two-species reaction model are solved using a hybrid sixth-order Weighted Essentially Non-Oscillatory-Centered Difference scheme combined with a structured adaptive mesh refinement technique. Results show that after the shutdown of the hot jet, the detonation wave is successfully stabilized quasi-steadily in the supersonic model combustor together with periodic fluctuations of the detonation front. The formation of the quasi-steady propagation of detonation in the model combustor is mainly due to the combined effects of (i) pressure oscillations generated in the cavity, which facilitate the detonation propagation, and (ii) lateral mass divergence brought by the expanding wall which can lead to detonation attenuation, and an unburned jet associated with large-scale vortices resulting from a Prandtl-Meyer expansion fan. This expansion fan is generated because of the expanding wall which can contribute to the detonation stabilization. It is found that for an incoming velocity lower than the Chapman–Jouguet value, a quasi-steady propagation of the detonation wave cannot be achieved. However, for incoming velocity higher than the Chapman–Jouguet value, a stabilization can be realized. This is effectively due to the formation of a periodic process including four stages of forward propagation, detonation attenuation, backward propagation and detonation bifurcation, indicating the influence of the supersonic model combustor on the overall process
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