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Understanding how protrusions, such as fins attached to flat or streamlined bodies, affect aerodynamics, especially in high-speed contexts, is vital for aerospace applications. These protrusions significantly influence overall aerodynamics and require a comprehensive understanding for accurate analysis and prediction of aerodynamic performance. This understanding is particularly critical in supersonic flight, where even minor aerodynamic disturbances can impact vehicle stability and efficiency. Therefore, a thorough understanding of protrusion-induced flow phenomena is essential for advancing aerospace engineering and improving supersonic vehicle performance and safety. The present paper focuses on the complex supersonic flow over a vertical fin, using a combination of experimental and computational methods. The study aims to understand how variations in fin height influence the behavior of the Lambda shock and any resulting changes in shock length. Specifically, the paper investigates different fin height-to-diameter (H/D) ratios ranging from 0.5 to 1.5 in steps of 0.25. To achieve this, both experimental testing in a supersonic wind tunnel and numerical simulations using the commercial CFD tool ANSYS-FLUENT are employed. Through this dual approach, the paper seeks insights into the characteristics of the Lambda shock and its effects on key aerodynamic parameters, such as shock strength and drag coefficient. By thoroughly investigating these aspects, the paper contributes to a deeper understanding of the complex flow phenomena associated with supersonic flow over vertical fins, potentially guiding the design and optimization of aerospace vehicles. The outcomes indicate that a fin height of 12 mm (H/D=1.0) provides the best balance in terms of pressure distribution, Lambda shock length, and drag coefficient, making it the optimal choice for enhancing aerodynamic stability and performance in supersonic conditions.

The Weighted Essentially Non-Oscillatory (WENO) scheme is a high-order accurate and computationally efficient numerical method widely used for simulating compressible fluid flows, particularly in scenarios involving shocks, discontinuities, and complex flow structures. This review paper provides a comprehensive overview of the WENO scheme, focusing on its application to compressible flow simulations. The paper introduces the fundamentals of the WENO scheme, including its core interpolation and reconstruction methodologies, and explores its ability to capture sharp gradients and complex flow phenomena, such as shock waves and turbulence. The review highlights advancements in WENO schemes, including their integration with finite volume and finite difference methods, and examines their performance in various compressible flow problems, including supersonic flows, shock-vortex interactions, and multiphase flows. Furthermore, the paper discusses the challenges of applying WENO to incompressible flows and weakly compressible regimes, emphasizing recent developments and potential adaptations. Current research trends, such as hybrid approaches and modified WENO schemes, are also reviewed, identifying opportunities for future exploration. This paper offers a critical evaluation of WENO schemes, serving as a valuable resource for researchers and practitioners in computational fluid dynamics.

The flow field in a two-dimensional hypersonic mixed-compression inlet in a freestream Mach numbers of M∞ =2.0, 3.0, and 5.0 are numerically solved to understand the effect of throat area variation. The exit area ratio variation is simulated by placing a plug insert at different axial locations at the exit of the model. The flow field is achieved computationally by solving the Reynolds Averaged Navier-Stokes equations in a finite volume framework. For each flow condition, the variation in shock structure is analyzed and the variation of the oblique shock wave angle with the mass flow rate is calculated theoretically and compared with the present CFD analysis. The variation in oblique shock angle is calculated in terms of the mass flow rate by considering the capture area and spillage flow through the inlet. The theoretical results suggest that the method can predict the inlet operating conditions at different freestream Mach numbers and area ratios. This method can quantify the reduction in mass flow rate due to the throttling effect by analyzing the flow field shock pattern. The effects of various important performance parameters such as free stream Mach number, total pressure recovery, and mass flow ratio were then numerically investigated. As the Mach number is increased, the total pressure recovery is reduced, but the maximum value of the mass flow rate is increased. The analysis is also focused on the effect of throat area variation on performance parameters at each Mach number. The characteristic curve of the inlet is then obtained for each free stream Mach number.

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.

Dynamic model of geometrical nonlinear deformations of functionally graded carbon nanotubes-reinforced composite cylindrical shell is obtained. Reddy higher-order shear deformation theory is used to derive this model. The finite-degree-of-freedom nonlinear system, which describes the structure nonlinear self-sustained vibrations, is obtained using the assumed-mode method. The linear piston theory is used to describe the supersonic flow. The loss of the cylindrical shell dynamic stability owing to the Hopf bifurcations is analyzed. The self-sustained vibrations, which describe the circumferential traveling waves flutter, occur due to this bifurcation. The harmonic balance method is applied to analyze these self-sustained vibrations. The properties of the circumferential traveling waves are analyzed.

Small-scale and supersonic convergent-divergent type micro-nozzles with characteristic sizes of around a few centimeters and exit and throat radii of tenths of millimeters were the subjects of this study. Using the schlieren Z-type optical technique, the supersonic airflows established at the exit of seven nozzles were visualized. The dependence of the shock cell characteristics on the nozzle pressure ratio (NPR), defined as the ratio of stagnation pressure to atmospheric pressure, was analyzed. The dependence of the nozzle thrust and the specific impulse on the NPR ratio and the mass flow rate was also studied using a simple device based on concepts of fluid mechanics. The results obtained are in agreement with similar results obtained in recently published research on double-bell nozzles. The thrust of all nozzles depends linearly on the shock-cell spacing, which is one of the most relevant findings of this research. In other words, the output airflow structure determines the performance of the nozzles, such as the thrust or the specific impulse they produce. These small nozzles offer significant advantages over conventional nozzles in low energy consumption and lower manufacturing cost, making them suitable for scientific research in space micro-propulsion and cooling microelectronic systems, among other applications.

The effect of instantaneous energy release (explosion) in the gas bubble region on supersonic flow past blunt bodies (sphere) and pointed bodies (ogival body and cone-cylinder combination) is considered when the explosion occurs in unperturbed freestream flow in the immediate neighborhood of the bow shock. Physically, such an effect on the flow can occur with energy input in the region of electric gas discharge or with detonation of a combustible gas mixture inside the bubble. It is found that, in addition to the direct effect of the explosive shock wave on the surface of the body, significant non-stationary changes in the gas-dynamic flow regimes past the bodies occur during the interaction of the bow shock with the dynamically varying explosion region (shock-compressed layer and cavity). In particular, focusing and cumulation effects, which can lead to secondary effects, are noted. The momentum of the latter is comparable to or even greater than the momentum of the direct impact of the blast wave.

Purpose The purpose of this paper is to demonstrate the aerodynamic behavior of a supersonic combustion test bench (SCTB) components, as the transition piece and the combustor of a scramjet (supersonic combustion ramjet), manufactured by 3D printing or additive manufacturing (AM). Design/methodology/approach For the dimensional and structural analysis of the manufactured models, a portable 3D scanner was used to generate the mesh of its dimensions, and to compare them before and after the experiments, a roughness measuring system was also used to verify the roughness inside the models before and after the tests, as roughness is an important parameter because it directly affects the boundary layer. For the visualization of the flow, the non-intrusive schlieren optical technique was used. Findings The experiments were carried out on the SCBT for Mach 2 flows, using the manufactured prototypes and showed that there was no structural and dimensional change of the model after the test batteries. It was found that the roughness presented by the material did not affect the quality of the flow generated. This shows that the investigated material can also be applied in experiments with supersonic flow. Originality/value This paper presents that it is possible to use in ground test facilities, for the studies of supersonic flow (in cold condition), pieces and models manufactured by 3D printing without affecting the quality of the flow generated during the experiments. This study presents a new perspective to approach AM applied in the studies of supersonic flows.

The two-dimensional electrogasdynamic problem of anomalous glow discharge on the surface of a sharp plate in supersonic flow of a perfect gas is solved using the system of Navier–Stokes equations to describe thermogasdynamic processes in the boundary layer and the two-temperature two-fluid diffusion-drift model of gas-discharge plasma to determine the electrodynamic structure of the discharge. The near-electrode regions of space charge and the external electrical circuit consisting of a power source and an ohmic resistance are taken into account. The influence of the magnetic field which is transverse to gas flow and has the induction of up to 0.03 T on the structure of boundary layer and glow discharge is studied. The electrogasdynamic structure of anomalous near-surface discharges is studied numerically over a wide range of gas flow velocities (M = 5–20), the free-stream pressures ( p = 0.6–5 Torr), the electrode voltages, and the electric currents through the discharges. The electrodynamic structure of the gas-plasma flow near the electrodes and the effect of the glow discharge on the pressure and temperature distributions along the surface of the plate are also studied.

The generation of high-quality computational mesh plays an important role in high-accuracy computation of supersonic flows over bodies. Priority is given to mesh adaptation to discontinuities, primarily, to bow shocks. In this paper, a mesh is treated as a mechanical system with elastic connections and an algorithm for mesh adaptation to a flow field with a bow shock is considered. The application of the algorithm to a typical structured mesh leads to mesh refinement in high field gradient regions achieved by drawing mesh lines into the shock vicinity, while maintaining the quality of the mesh elements. The problems considered demonstrate the possibility of applying the described algorithm to realistic flow problems of practical interest.