Search Results (93)

A combined flow control method, integrating nanosecond pulsed surface dielectric barrier discharge (NS-SDBD) with pulsed jets, is proposed to address the challenge of low mixing efficiency in supersonic combustion. Numerical validation and mechanism analysis were conducted by solving the three-dimensional unsteady Reynolds-averaged Navier–Stokes (RANS) equations, coupled with the shear stress transport (SST) k–ω turbulence model. The simulations were carried out under a Mach 2.8 inflow condition with a 50 kHz pulsed frequency for H₂ jets. The results demonstrate that, compared to the steady jet case, the combined control scheme increases the combustion product mass flow rate by 27.1% and enhances combustion efficiency by 26.8%. The average temperature in the wake region increases by 65 K, while the total pressure recovery coefficient shows only a marginal change. The pressure disturbance center evolves along the outer edge of the counter-rotating vortex pair (CVP) and is eventually absorbed by the vortex core. This process generates favorable velocity and vorticity perturbations, which enhance O₂ entrainment into the CVP and increase the average wake temperature. Meanwhile, the strengthened reflected shock induces favorable velocity perturbations in the upper shear layer of the wake and further elevates the local temperature. Full article

Given the undoubtable similarities between the dynamic behavior of the vehicular traffic flow in terms of its response to boundary condition alterations dictated in the form of obstacles, and the specific case of supersonic compressible fluid flow fields, the current manuscript addresses developing a target trajectory for the overtaking maneuver of autonomous vehicles. The path-planning is pursued in analogy to the governing principles of the supersonic compressible fluid flow fields, with the specific definition of a physically meaningful dimensionless group, namely the Traffic Mach number (M_T), which grants the initial access point to the said set of fundamental equations. This practical application is a follow-up to the primarily established proof-of-concept level introduction and analysis of the more general case of collision avoidance for autonomously driven vehicles in accordance with the supersonic compressible fluid flow field, where the Traffic Mach number was first introduced. The proposed trajectory is then taken to the next block of the investigation, namely the tracking and control aspects of the maneuvering vehicle’s dynamics. The path tracking controller is designed based on sliding mode control technique and the algorithm is applied on a 7-DOF simulation model, used for validation and discussion of results. The proposed method is shown to be suitable for overtaking maneuvers of autonomous vehicles, whilst meeting the criteria for a relative velocity from the constant-velocity vehicle ahead of the road in the supersonic regime based on the defined Traffic Mach number. The results are then presented, first, in the scope of the aerodynamics field configuration and their verifications, followed by the vehicle dynamics remarks showing the practicality of the proposed method in terms of vehicle motion. It is observed that the distance corresponding to the delayed maneuver maximizes at highest velocities of the ego vehicle, consistent with the highest M_T values, yet in all simulated cases, the control system of the vehicle model was capable of performing the maneuver based on the assigned trajectories through the present model. Full article

Flow quality at the engine face, especially total pressure recovery and swirl, is central to the performance and stability of external compression supersonic inlets. Steady-state RANS-based numerical computations are performed to quantify bleed/swirl trade-offs in a single-ramp intake. The CFD simulations were performed first without a bleed system over M_∞ = 1.4–1.9 to locate the practical onset of a bleed requirement. The deterioration in pressure recovery and swirl beyond M_∞ ≈ 1.6, which is consistent with a pre-shock strength near the turbulent separation threshold, motivated the use of a bleed system. The comparisons with and without the bleed system were performed next at M_∞ = 1.6, 1.8, and 1.9 across the operation map parameterized by the flow ratio. The CFD simulations were performed using ANSYS Fluent, with a pressure-based coupled solver with a realizable k-ε turbulence model and enhanced wall treatment. The results provide engine-face distortion metrics using a standardized ring to sector swirl ratio alongside pressure recovery. The results show that bleed removes low-momentum near-wall fluid and stabilizes the terminal–shock interaction, raising pressure recovery and lowering peak swirl and swirl intensity across the map, while extending the stable operating range to a lower flow ratio at a fixed M_∞. The analysis delivers a design-oriented linkage between shock/boundary-layer interaction control and swirl: when bleed is applied at and above M_∞ = 1.6, the separation footprints shrink and the organized swirl sectors weaken, yielding improved operability with modest bleed fractions. Full article

There have been many attempts to model the flow of vehicular traffic in analogy to the flow of fluids. Given the evident change in distance between vehicles driving in platoons, the compressibility of traffic flow is inferred and, considering the reaction time-scales of the driver (human or autonomous), it is argued that this compressibility is increased as relative velocities increase—giving the lag in imposed redirection by the driver and the controller units a higher relative importance. Therefore, a supersonic compressible flow field has been opted for as the most analogous base flow. On this point, added to by the overall extreme similarities of the two above-mentioned flows, the non-dimensional group of the traffic Mach number M_T has been defined in the present research, providing the possibility of calculating a suggested flow field and its corresponding shockwave systems, for any given obstacle ahead of the traffic flow. This suggested flow field is then taken as the basis to obtain trajectories designed for avoiding collision with the obstacle, and in compliance with the physics of the underlying analogous fluid flow phenomena, namely the internal supersonic compressible flow around a double wedge. It should be noted that herein we do not model the traffic flow but propose these trajectories for more optimal collision avoidance, and therefore the above-mentioned similarities (explained in detail in the manuscript) suffice, without the need to rely on full analogies between the two flows. The manuscript further analyzes the applicability of the proposed analogy in the path-planning process for an autonomous passenger vehicle, through dynamics and control of a full-planar vehicle model with an autonomous path-tracking controller. Simulations are performed using realistic vehicle parameters and the results show that the fluid flow analogy is compatible with the vehicle dynamics, as it is able to follow the target path generated by fluid flow calculations with minor deviations. Simulation results demonstrate that the proposed method produces smooth and dynamically consistent trajectories that remain stable under varying traffic scenarios. The controller achieves accurate path tracking and rapid convergence, confirming the feasibility of the fluid-flow analogy for real-time vehicle control. Full article

The increasing demand for effective flow control in supersonic boundary layers, particularly for mitigating shock-wave boundary-layer interactions, underscores the need to explore optimized micro-vortex generator (MVG) configurations. This study investigates the aerodynamic performance of two different MVG configurations: a two-MVG setup with a pair of close parallel-positioned MVGs and a three-MVG arrangement that includes an additional upstream unit. Both are examined within a Mach 2.5 flow regime, aiming to improve mixing and energize the boundary layer. Large Eddy Simulations (LES) were performed using high-order numerical schemes. A newly developed vortex identification method was utilized to characterize vortex structures, while turbulent kinetic energy (TKE) metrics were integrated to quantify turbulence. Findings reveal that the two-MVG configuration produces regular, symmetric vortex pairs with limited interaction. This results in a steady increase in TKE and a thickened momentum boundary layer—indicative of notable energy loss. In contrast, the three-MVG setup generates more intricate and interactive vortex formations that significantly elevate TKE levels, rapidly expand the turbulent region, and reduce energy loss downstream. The peak TKE occurs before tapering slightly. Instantaneous flow analysis further highlights chaotic, hairpin-dominated vortex structures in the three-MVG case, compared to the more orderly ones observed in the two-MVG case. Overall, the three-MVG configuration demonstrates superior mixing and boundary-layer energization potential, albeit with greater structural complexity. Full article

The interaction between a shock wave and a boundary layer promotes corner separation and prevents performance enhancement in a supersonic compressor cascade. Different vortex generator (VG) designs are presented to control corner separation in a supersonic compressor cascade, including endwall VGs (EVG), suction surface VGs (SVG), and combined endwall and suction surface VGs (E-SVGs). It is demonstrated that EVG and coupled E-SVGs reduce losses in the supersonic compressor cascade. For an optimal EVG, the total loss is reduced by 24.6% and the endwall loss is reduced by 33.6%. The coupled E-SVG better controls corner separation and reduces endwall losses by 56.9%. The suppression mechanism is that vortices alter the direction of the separated flow, allowing it to overcome the chordwise pressure gradient. Moreover, the VGs change the shock structure near the endwall. For the EVG, clockwise vortices are effective in controlling corner separation due to their minor effect on the shock structure near the endwall. However, anticlockwise vortices are not suitable for controlling corner separation in the supersonic compressor because they increase the shock strength induced by the VG. The control mechanism of the coupled E-SVG on corner separation is also discussed. Full article

Computational fluid dynamics (CFD) has aided the development, design, and analysis of hypersonic airbreathing propulsion technologies, such as scramjets. The complex flow field in a scramjet isolator has been the subject of intense interest and study for several decades. Many features of this flow field also occur in supersonic wind-tunnel nozzles and diffusers. Computational analysis of these topics has frequently provided immense insight into the actual functionality and performance. Research presented in this work supports scientific investigation and understanding of a less-researched topic, which is shock–shock interference and interaction with the boundary layer in supersonic internal flows, as well as the passive control of its adverse effects to prevent the onset of unstart in a scramjet isolator. This computational investigation is conducted on a backpressured isolator and a modified three-dimensional shock-tube to represent a scramjet isolator with ram effects provided by high-pressure gas and high-speed flow provided by a supersonic inflow. Computational results for the backpressured isolator have been validated against available measured time-averaged wall pressure data. The modified shock-tube provided an opportunity to study the shock–shock interference and shock–boundary-layer interaction effects that would occur in a scramjet isolator or a ram-accelerator when the high-speed flow from the inlet interacted with the shock produced due to the combustor pressure traveling and meeting in the isolator. An assessment of wall cooling effects on these phenomena is presented for both the backpressured isolator and the modified shock-tube. Full article

This study investigates supersonic flow within a nozzle under low-pressure conditions at the continuum mechanics boundary. This phenomenon is commonly encountered in applications such as the differentially pumped chamber of an Environmental Scanning Electron Microscope (ESEM), which employs an aperture to separate two regions with a great pressure gradient. The nozzle geometry and flow control in this region can significantly influence the scattering and loss of the primary electron beam traversing the differentially pumped chamber and aperture. To this end, an experimental chamber was designed to explore aspects of this low-pressure regime, characterized by a varying ratio of inertial to viscous forces. The initial experimental results obtained using pressure sensors from the fabricated experimental chamber were utilized to refine the Ansys Fluent simulation setup, and in this combined approach, initial analyses of supersonic flow and shock waves in low-pressure environments were conducted. The refined Ansys Fluent system demonstrated a very good correspondence with the experimental findings. Subsequently, an analysis of the influence of surface roughness on the resulting flow behavior in low-pressure conditions was performed on this refined model using the refined CFD model. Based on the obtained results, a comparison of the influence of nozzle roughness on the resulting electron beam scattering was conducted for selected low-pressure variants relevant to the operational conditions of the Environmental Scanning Electron Microscope (ESEM). The influence of roughness at elevated working pressures within the ESEM operating regime on reduced electron beam scattering has been demonstrated. At lower pressure values within the ESEM operating regime, this influence is significantly diminished. Full article

Shockwaves induce considerable flow separation loss; it is essential to reduce this using the flow control method. In this manuscript, a method for suppressing flow separation in turbomachinery through a constant adverse-pressure gradient was investigated. The first-passage shock was split into a compression wave system of the vane suction surface. The aim of this was to reduce loss from shockwave/boundary layer interactions (SWBLIs). This method promotes the performance parameters of the supersonic compressor cascade. The investigation targets were a baseline cascade and the improved system. Both cascades were numerically studied with the aid of the Reynolds-averaged Navier–Stokes (RANS) method. The simulation results of the baseline cascade were also validated through experimentation, and a further physical flow analysis of the two cascades was conducted. The results show that the first-passage shockwave was a foot above the initial suction surface, with a weaker incident shock along with a clustering of the compression wave corresponding to the modified cascade. It was also concluded that the first-passage shockwave foot of the baseline cascade was replaced with a weak incident shock, and a series of compression waves emanated from the adopted negative-curvature profile. The shock-induced boundary layer separation bubble disappeared, and much smaller boundary layer shape factors over the SWBLI region were obtained for the improved cascade compared to the baseline cascade. This improvement led to a high level of stability in the boundary layer state. Sensitivity analyses were performed through different simulations on both cascades, unveiling that the loss in total pressure was lower in the case of the updated cascade as compared to the baseline. Full article

Micro-vortex generators (MVGs) are widely utilized as passive devices to control flow separation in supersonic boundary layers by generating ring-like vortices that mitigate shock-induced effects. This study employs large eddy simulation (LES) to investigate the flow structures in a supersonic boundary layer (Mach 2.5, Re = 5760) controlled by two MVGs installed in tandem, with spacings varying from 11.75 h to 18.75 h (h = MVG height), alongside a single-MVG reference case. A fifth-order WENO scheme and third-order TVD Runge–Kutta method were used to solve the unfiltered Navier–Stokes equations, with the Liutex method applied to visualize vortex structures. Results reveal that tandem MVGs produce complex vortex interactions, with spanwise and streamwise vortices merging extensively, leading to a significant reduction in vortex intensity due to mutual cancellation. A momentum deficit forms behind the second MVG, weakening that from the first, while the boundary layer energy thickness doubles compared to the single-MVG case, indicating increased energy loss. Streamwise vorticity distributions and instantaneous streamlines highlight intensified interactions with closer spacings, yet this complexity diminishes overall flow control effectiveness. Contrary to expectations, the tandem configuration does not enhance boundary layer control but instead weakens it, as evidenced by reduced vortex strength and amplified energy dissipation. These findings underscore a critical trade-off in tandem MVG deployment, suggesting that while vortex interactions enrich flow complexity, they may compromise the intended control benefits in supersonic flows, with implications for optimizing MVG arrangements in practical applications. Full article

Current research on the transition mechanisms induced by moderate-height roughness elements remains insufficiently explored. Hence, direct numerical simulation (DNS) and BiGlobal stability analysis are employed in this study to investigate boundary layer transition from laminar to turbulent flow induced by moderate-height isolated roughness elements and roughness strips under a supersonic freestream at Mach 3.5. Analysis of DNS results reveals that the isolated roughness element induces transition within the boundary layer, characterized by two high-speed streaks in the wake. This transition is attributed to the coupling between the separated shear layer at the roughness apex and the downstream counter-rotating vortex pair (CVP). BiGlobal stability analysis further identifies that symmetric eigenmodes dominate the transition process in the wake, actively promoting flow destabilization. Conversely, the roughness strip configuration suppresses transition, with only attenuated high-speed streaks persisting in the near wake before complete dissipation. The wake flow exhibits multiple CVPs and adjacent horseshoe vortex pairs interacting with the shear layer, with antisymmetric modes dominating this process. These findings provide technical foundations and theoretical frameworks for predicting and controlling roughness-induced transition. Full article

The thermal vibration of thick Terfenol-D control law on functionally graded material (FGM) plates/cylindrical shells in nonlinear unsteady supersonic flow with third-order shear deformation theory (TSDT) is investigated by using the generalized differential quadrature (GDQ) method. The effects of the coefficient term of TSDT displacement models on the thermal stress and center displacement of Terfenol-D control law on FGM plates/cylindrical shells in nonlinear unsteady supersonic flow are investigated. The coefficient term of TSDT models of thick Terfenol-D control law on FGM plates/cylindrical shells provide an additional effect on the values of displacements and stresses. Full article

This work studies the influence of a growing boundary layer on the process of supersonic flow around an aerodynamic body. The task is to select and implement in an experiment the parameters of a supersonic flow and to study the flow pattern near the surface of an aerodynamic body at different viscosity values for the incoming flow. Visualization of the shock wave configuration in front of the body and studying the change in the pressure field in the flow region under these conditions is the main goal of this work. The experiment was carried out on an experimental stand created on the basis of a shock tube. The aerodynamic body under study (a semi-cylinder pointed along a circle or an ellipse) was placed in a supersonic nozzle. The model was clamped by lateral transparent walls, which were simultaneously a source of boundary layer growth and the viewing windows for visualizing the flow. For selected modes with Reynolds numbers from 8200 to 45,000, schlieren flow patterns and pressure distribution fields near the surface of the streamlined models and the plate of the growing boundary layer were obtained. The data show a complex, unsteady flow pattern realized near the model which was caused by the viscous-inviscid interaction of the boundary layer with the bow shock wave near the wall. Full article

This work explores a potential methodology for rectangular jet noise reduction that employs nozzle unsteady microjet excitation. Using high-fidelity computational studies and spectral analyses, major jet noise sources impacted by the applied actuation are identified. A heated supersonic rectangular jet is considered with a nozzle aspect ratio of 2:1 at a Mach number of 1.5. The current study essentially validates the hypothesis of a previous reduced-order analysis that predicted jet noise reduction through jet excitation at the harmonic or subharmonic of the dominant frequency associated with jets’ large-scale structures. Such noise reduction was attributed to the excitation-induced nonlinear energy exchange between the coherent modes. In the current study, the synthetic microjet actuation of the jet plume shear layer using $1\%$ of the jet mass flow rate is implemented at the excitation ports located at the nozzle lip and directed along the jet axis. A resulting jet noise reduction of up to 4 dB at the peak radiation angle is predicted. An analysis of the near-field Spectral Proper Orthogonal Decomposition (SPOD) results provides further insights into the impact of jet actuation on the modification of jet flow structures, thus addressing the effectiveness of the proposed noise control methodology. Full article

Numerical investigations were conducted to analyze the coupling effects of pulsed H₂ jets and nanosecond-pulsed actuation (NS-SDBD) in a supersonic crossflow. The FVM was employed to solve the multi-component 2D URANS equations with the SST k-omega turbulence model, while H₂-air combustion was described using a seven species–seven reactions chain reaction model, and the plasma thermal effect was represented by a phenomenological model. The backward-facing step flows with an inlet Mach number of 2.5 and a pulsed jet frequency of 10 kHz under different actuation conditions were simulated. The combustion enhancement mechanism under an actuation frequency of 20 kHz was analyzed. Research indicates that compression waves induced by NS-SDBD enhance H₂-air mixing and facilitate temperature transport as the flow progresses. This progress is significantly associated with the flow structures generated by pulsed jets. Under this condition, the fuel utilization rate in the flow field increased by 61.2%, the total pressure recovery coefficient increased by 5.34%, and the outlet total temperature slightly increased even with a 50% reduction in fuel flow rate. Comparative analysis of different actuation cases demonstrates that evenly distributed actuation within the jet cycle yields better effects. The innovation of this study lies in proposing and exploring a potential method to address inadequate combustion under high-speed inflow conditions, which couples NS-SDBD with pulsed hydrogen jets. Full article