Search Results (21)

The design of heat and drag reduction systems for hypersonic vehicles has garnered widespread global attention. In this study, the Navier–Stokes equations and the SST k-ω turbulence model are employed to establish a simulation model for heat and drag reduction induced by a forward-facing cavity. The numerical methods are validated using existing experimental results. The oscillation characteristics of the bow shock wave at the head and the shock inside the cavity in hypersonic flows are investigated. The heat and drag reduction mechanisms of the forward-facing cavity are discussed. The effects of the diameter and depth of the cavity on drag and heat reduction are comprehensively analyzed. The obtained results show that a reduction in drag and heat is achieved when a forward-facing cavity is added to the vehicle. The main reasons for this heat reduction are the cold ring mechanism and the energy conversion mechanism. The size of the cold ring is significantly affected by the cavity diameter, whereas the energy conversion mechanism is more sensitive to variations in diameter. The maximum reduction in heat load is 2.2%, and the maximum reduction in the Stanton number is 25.3%. Increases in both diameter and depth enhance drag reduction, achieving an average drag reduction of approximately 1.65%. 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

The main objective of this study is to demonstrate the active influence on the location of the bow shock wave, as well as on the parameters of an aerodynamic body, of a gas discharge organized near the frontal surface, between the body and the bow shock wave. The research is carried out using both experimental and numerical methods at the freestream Mach number M = 6.8. The working gas is xenon. It is shown that the location of the steady bow shock wave, along with the current and power of the discharge, is associated with the change in the adiabatic index of the plasma created by the discharge, which, in turn, is determined by plasma parameters such as the degrees of nonequilibrium and the degree of ionization. It is shown that the adiabatic index with the power supplied to the impact zone in the range of 30–120 kW can both increase and decrease in the range of 1.25–1.288. A study of the discharge-created plasma zone is conducted, and the correspondence between the gas discharge current and power and the average parameters in the plasma zone created by the discharge are presented. A good agreement between the numerical and experimental data is shown. The results obtained can be useful in the development of control systems for high-speed flows based not only on the effects of heating but also on the impact of plasma parameters. Full article

This review examines ultra-low frequency (ULF) waves across different planetary environments, focusing on Earth, Mercury, and Saturn. Data from spacecraft missions (CHAMP, Swarm, and Oersted for Earth; MESSENGER for Mercury; and Cassini for Saturn) provide insights into ULF wave dynamics. At Earth, compressional ULF waves, particularly Pc3 waves, show significant power near the equator and peak around Magnetic Local Time (MLT) = 11. These waves interact complexly with Alfvén waves, impacting ionospheric responses and geomagnetic field line resonances. At Mercury, ULF waves transition from circular to linear polarization, indicating resonant interactions influenced by compressional components. MESSENGER data reveal a lower occurrence rate of ULF waves in Mercury’s foreshock compared to Earth’s, attributed to reduced backstreaming protons and lower solar wind Alfvénic Mach numbers, as ULF wave activity increases with heliocentric distance. Short Large-Amplitude Magnetic Structures (SLAMS) observed at Mercury and Saturn show distinct characteristics compared to those of Earth, including the presence of whistler precursos waves. However, due to the large differences in heliospheric distances, SLAMS (their temporal scale size correlate with the ULF wave frequency) at Mercury are significantly shorter in duration than at Earth or Saturn, since the ULF wave frequency primarily depends on the strength of the interplanetary magnetic field. This review highlights the variability of ULF waves and SLAMS across planetary environments, emphasizing Earth’s well-understood ionospheric interactions and the unique behaviours observed for Mercury and Saturn. These findings enhance our understanding of space plasma dynamics and underline the need for further research regarding planetary magnetospheres. Full article

During the operation of hypersonic vehicles, a reciprocal coupling effect is manifested between the inlet and the combustion chamber. This results in an unavoidable non-uniformity of conditions at the combustion chamber’s entrance, which, in turn, influences the fuel mixing within the chamber. The present study employed the Reynolds-averaged Navier–Stokes (RANS) equations to perform a numerical simulation of an X-51-like vehicle, with a focus on examining the impact of isolation section length and multi-injection strategies on the fuel mixing characteristics within the combustion chamber under conditions of non-uniform inflow. The findings indicated that a supersonic non-uniform inlet triggers incident shock waves, leading to a non-uniform pressure distribution across the flow section. Moreover, the position of injection was found to be pivotal in regulating penetration depth and mixing efficiency. The incident shock wave, bow shock, and boundary layer separation shock interacted with each other to increase local pressure. The coupling of high and low pressures generated an adverse pressure gradient that led to boundary layer separation, which further enhanced fuel penetration depth. Full article

This review examines studies aimed at the organization of energy (non-mechanical) control of high-speed flow/flight using spatially multi-component plasma structures and combined energy deposition. The review covers selected works on the experimental acquisition and numerical modeling of multi-component plasma structures and the use of sets of actuators based on plasma of such a spatial type for the purposes of control of shock wave/bow shock wave–energy source interaction, as well as control of shock wave–boundary layer interaction. A series of works on repetitive multiple laser pulse plasma structures is also analyzed from the point of view of examining shock wave/bow shock wave–boundary layer interaction. Self-sustained theoretical models for laser dual-pulse, multi-mode laser pulses, and self-sustained glow discharge are also considered. Separate sections are devoted to high-speed flow control using combined physical phenomena and numerical prediction of flow control possibilities using thermal longitudinally layered plasma structures. The wide possibilities for organization and applying spatially multi-component structured plasma for the purposes of high-speed flow control are demonstrated. Full article

Aluminum alloy 6082-T6 is an important material for manufacturing the outer skin of high-speed trains, and laser shock forming can realize the rapid forming of complex-shaped plates. In order to improve the efficiency of the simulation modeling of laser shock forming for aluminum alloy 6082-T6, Python scripting language was used for the secondary development of Abaqus. A plugin was utilized to simulate and analyze the laser shock forming process of aluminum alloy 6082-T6. The coordinates of the plate after laser impact molding were measured using a coordinate measuring machine to calculate the arc bow height of the plate. The accuracy of the simulation model was verified by comparing with the simulation results. The deformation characteristics of plastic strain and arc height of aluminum alloy 6082-T6 under different laser process parameters were analyzed. The simulation plugin has a concise interface, high operability, and accurate results with the other parameters unchanged. When the laser energy is 5 J, 6 J, and 7 J, the corresponding arc heights are 5.9 mm, 6.6 mm, and 7.2 mm, respectively. As the thickness of the sheet increases, the deformation changes from concave at 1 mm to convex at 2 mm, 3 mm, 4 mm, and 5 mm. As the spot size increases from 1 mm to 5 mm, the transmission mode of the shock wave gradually changes from spherical wave to planar wave, and the arc height of the sheet increases from 4.6 mm to 8.2 mm. With the increase in the spot overlap rate, the impact area accumulates residual stress, and the arc height of the sheet is 5.7 mm, 6.6 mm, 7.3 mm, and 8.5 mm, respectively. The secondary development of ABAQUS 2021 using Python 3.6 scripting language has improved the efficiency of simulation modeling and provided reference for rapidly predicting the deformation characteristics of aluminum alloy 6082-T6 under different laser process parameters. Full article

The goal of this paper is to investigate the aerodynamic and aerothermodynamic behavior of the Schiaparelli capsule after the deployment of a supersonic disk-gap-band (DGB) parachute during its re-entry phase into the Martian atmosphere. The novelty of this work lies in the investigation by LES (large-eddy simulations) of the coupled interaction of the flow field generated behind the capsule and that in front of the flexible DGB parachute. These simulations are performed at an altitude of 10 km and a Mach number around 2, i.e., a regime in which large canopy-area oscillations are observed. LES results have shown a strong interaction between the bow shock, the recompression and expansion waves, high pressure, density and temperature gradients, heat flux towards the airstream and the body implying turbulence generation, ingestion, and amplification through the shock waves. Vortices released from the capsule at a frequency of about 52 Hz and 159 Hz, corresponding to Strouhal numbers of ~0.2 and 0.75, respectively, are the main factors responsible for the instabilities of the hypersonic re-entry capsule and the disk-gap-band parachute coupled system. The nonlinear turbulence flow field generated at the capsule back is amplified when passing the parachute bow shock, and this is responsible for the non-axisymmetric behavior around and behind the parachute that caused the uncontrolled capsule oscillations and the Schiaparelli mission failure. In fact, an LES of the parachute without the capsule, for the same conditions, show a completely axisymmetric field, varying in time, but axisymmetric. In order to avoid this turbulence amplification, dampening of the vortex shedding is critical. Different techniques have been already proposed for other applications. In the case of capsule re-entry, due to the high temperatures in front of the capsule behind the bow shock since air plasma is generated, damping of the vortex shedding could be achieved by means of magnetohydrodynamic (MHD) control. Full article

In the combustion chamber of scramjets, fuel jets interact with supersonic airflow in the form of a liquid jet in crossflow (LJIC). It is difficult to achieve adequate jet–crossflow mixing and the efficient combustion of fuel in an instant. Large eddy simulation (LES), the coupled level-set and volume of fluid (CLSVOF) method, and an adaptive mesh refinement (AMR) framework are used to simulate supersonic LJICs in this article. This way, LJIC atomization characteristics and mechanisms can be further explored and analyzed in detail. It is found that the surface waves of the liquid column exist in a two-dimensional form, including vertical and spanwise directions. Column breakup occurs when all the spanwise surface waves between adjacent vertical surface waves break up. Bow shock waves, composed of multiple connected arcuate shock waves, are dynamic and will change with the evolution of the liquid column. The vortex ring movement of supersonic LJICs, whose trends in the vertical and spanwise directions are different, is relatively complex, which is due to the complex and time-dependent shape of liquid columns. Full article

The problem of noise reduction in supersonic aircraft design is one of the key problems, the solution of which largely determines the speed of development of supersonic aviation as a whole. The present study examines the noise generation during flights of supersonic civil aircraft. The effect of a thermally stratified energy source (TSS) used to control the supersonic flow past a pointed cylinder aerodynamic model on the near-field and ground pressure signatures, as well as on the perceived loudness in decibels (PLdB) on the ground, is evaluated. The complex conservative difference schemes, Tomas’ waveform parameter method, and Stevens’ algorithm Mark VII are used for near-field modeling, obtaining the ground pressure signature, and the evaluation of the PLdB on the ground, accordingly. The fields of flow parameters and the dynamics of a drag force are researched at the variation of temperatures in layers of TSS and for different numbers of layers. Simulations showed that changing the surface pressure due to drag reduction does not necessarily imply a change in the PLdB on the ground. In particular, it has been shown that when performing the flow control at freestream Mach numbers 1.5–2 using TSSs with the number of layers from 2.5 to 7.5 and rarefaction parameters in the layers from 0.15 to 0.3, some weakening of the bow shock wave in the near-field pressure signature due to the effect of TSS occurs, and no additional noise impact on the ground is introduced. Full article

A simple and efficient flow field visualization method (based on shadow imaging) was applied in a direct-connect test to explore the influence of the momentum flux ratio and the jet angle on the formation and evolution of nitrogen jets in supersonic combustion chambers. The test setup adopts a rectangular flow passage to simulate a flight condition with Mach number of 6 and altitude of 25 km. The experimental results showed that (a) the flow field visualization method adopted in this paper can clearly register the formation and evolution of the shock wave structure in the flow field and the windward shear vortex on the jet surface. (b) The evolution process of the windward shear vortex is significantly affected by the jet angle. In particular, the stretching position of the windward shear vortex changed when the jet angle was obtuse. (c) The bow shocks showed local distortion due to the periodic generation of large-scale shear vortexes. (d) Under the working conditions of the test, the largest instability of the flow field was found for a jet angle of 120°. This work provides, on one hand, the experimental basis for clarifying the formation and evolution mechanism of transverse gas jets, and on the other, valuable data that can be used to validate numerical simulations. Full article

A new method of high-speed flow control using permanently operating thermally stratified energy deposition is presented. The paper focuses on the analysis of the dependence of the characteristics of a steady supersonic flow and an aerodynamic (AD) body on the temperature values in the layers of a stratified source and the possibility of making the transition from one steady flow mode to another by changing the temperature in the layers. A detailed visualization of the dynamics of the fields of density, pressure, temperature, and local Mach number is presented during the controlled establishment of steady flow modes. Multiple generation of the Richtmyer–Meshkov instability is shown. The sharp peaks accompanying the development of the Richtmyer–Meshkov instabilities were obtained, which remain in the steady flow mode established under the action of a stratified energy source. Basic approaches for controlling the bow shock wave, drag and lift (pitch) forces (at zero angle of attack), and the stability in a steady supersonic flow past an AD body using permanently operating thermally stratified energy source were developed. The possibility of initiating and damping self-sustained flow pulsations as well as the formation of a steady flow with oppositely directed constantly acting lift forces due to temperature changes in the layers of a thermally stratified energy source is shown. Full article

This study focused on the development of the unsteady impact of a thermally stratified energy source on a supersonic flow around an aerodynamic (AD) body in a viscous heat-conducting gas (air). Research was based on the Navier-Stokes equations. The freestream Mach number was 2. A new multi-vortex mechanism of the impact of a time-limited stratified energy source on the aerodynamic characteristics of a body was described. Almost complete destruction of the bow shock wave in the density field, due to the multiple generation of Richtmyer-Meshkov instabilities in the region of a stratified energy source, was obtained. The dependences of the dynamics of frontal drag and lift forces of a streamlined body on temperature in the source layers were studied. It was determined that, by changing the temperature in the layers of a stratified energy source, it was possible to obtain more intense vortices accompanying the Richtmyer-Meshkov instabilities, causing a temporary decrease in the drag force of an AD body and ensuring the emergence and unsteady change in the magnitude of the lift (pitch) forces. The main principles of unsteady flow control using a stratified energy source were established. Full article

This paper demonstrates the possibility of active magnetohydrodynamic (MHD) control of supersonic flows containing shock waves. The shock wave configurations that occur at the inlet to a supersonic diffuser and in front of a streamlined semicylindrical model are used for the purpose of investigation. The impact is carried out by organizing local gas discharge regions when applying a magnetic field transverse to gas discharge currents. It has been shown that by changing the local region of application, the intensity and the direction of the gas discharge currents, it is possible to change the intensity and direction of the ponderomotive force acting on the gas flow during MHD interaction. The ponderomotive force control allows for acting locally on the shape and position of shock waves, the speed and direction of the flow, and the increase or reduction of pressure near the surface of the streamlined body. The experiments were carried out on a gas dynamic setup based on a shock tube in a gas dynamic path, capable of creating supersonic flows in a wide range of Mach numbers at M = 4–7. There was a possibility of organizing the electric and pulsed magnetic fields with an intensity of up to 1.5 T. The given experimental Schlieren flow patterns and the analysis of the obtained data demonstrate the MHD effect on: the change in the angle of inclination of the attached shocks, both into increase and decrease; the bow shock wave approaching the body or the removal from it; and the change in the aerodynamic drag and lift force of the streamlined bodies. Full article

The paper presents two methods to improve the efficiency of supersonic flow simulation using arbitrarily shaped unstructured grids. The first method promotes increasing the numerical solution convergence rate and is based on the geometric multigrid method for initialization of the flow field. The method is used to obtain the initial field of distributed physical quantity values, which maximally corresponds to the converged solution. For this purpose, the problem simulation is performed on a series of coarse grids beginning from the coarsest one in this series. Upon completion of simulations, the solution obtained is interpolated to a finer grid and used for initialization of simulations on this grid. The second method allows increasing the numerical solution accuracy and is based on statically adapting the computational grid to the flow specifics. The static adaptation algorithm provides automatic refinement of the computational grid in the region of specific features of flow, such as shock waves typical for supersonic flows. This algorithm provides a better description of the shock-wave front owing to the local grid refinement, with the local refinement region being automatically selected. Results of using these methods are demonstrated for the two supersonic aerodynamics problems: the simulation of the bow shock strength at a given distance under axially symmetric body Seeb-ALR and a mock-up aircraft Lockheed Martin 1021. It is shown that in both cases, the numerical solution convergence rate is increased owing to the use of the geometric multigrid method for initialization and a higher quality and a higher accuracy of solution is gained owing to the local grid refinement (using static adaptation means) near the shock-wave front. Full article