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The nonlinear flutter of viscoelastic three-layer plates and shallow shells with a structure asymmetrical in thickness, flown around by a supersonic gas flow is studied in this paper. Mathematical models of problems on the flutter of viscoelastic three-layer plates, cylindrical panels, and shells with a structure asymmetrical in thickness, flown around by a supersonic gas flow are developed. The flutter of viscoelastic three-layer plates is studied in linear and nonlinear formulations. The critical flutter velocities of elongated plates are compared with the results obtained in previously published studies, where the solutions were obtained in an elastic formulation. The flutter of viscoelastic three-layer plates, cylindrical panels with a rigid filler that resists transverse shear, flown from the outside by a supersonic flow, was studied. It is shown that an increase in the geometric parameter characterizing the flexural rigidity of the bearing layers of three-layer structures leads to an increase in the flutter velocity by 25–40%.

A finite-difference lattice Boltzmann (LB) algorithm is described on general body-fitted coordinate systems. An alternative treatment for the implicit collision term of the Boltzmann–Bhatnagar–Gross–Krook equation is used, which completely removes the implicitness of the numerical scheme through using the characteristic of collision invariants. LB simulations are carried out for a two-dimensional supersonic viscous flow past a circular cylinder and the natural convection heat transfer in a circular enclosure with an inner hexagonal cylinder for the first time. The pressure coefficient distribution along the surface of the circular cylinder and the Nusselt number in the natural convection obtained from the simulations agree well with previous experimental measurements and/or classical computational fluid dynamics simulations. Comparisons of detailed flow patterns with other studies are also satisfactory.

Attempts to simulate compressible flows with moderate Mach number to relatively high ones using Lattice Boltzmann Method (LBM) have been made by numerous researchers in the recent decade. The stability of the LBM is a challenging problem in the simulation of compressible flows with different types of embedded discontinuities. The present study proposes an approach for simulation of inviscid flows by a compressible LB model in order to enhance the robustness using a combination of Essentially NonOscillatory (ENO) scheme and Shock-Detecting Sensor (SDS) procedure. A sensor is introduced with adjustable parameters which is active near the discontinuities and affects less on smooth regions. The validity of the improved model to capture shocks and to resolve contact discontinuity and rarefaction waves in the well-known benchmarks such as, Riemann problem, and shock reflection is investigated. In addition, the problem of supersonic flow in a channel with ramp is simulated using a skewed rectangular grid generated by an algebraic grid generation method. The numerical results are compared with analytical ones and those obtained by solving the original model. The numerical results show that the presented scheme is capable of generating more robust solutions in the simulation of compressible flows and is almost free of oscillations for high Mach numbers. Good agreements are obtained for all problems.

Heat and drag reduction on the nose cone is a significant issue for increasing the speed of the supersonic vehicles. In this paper, computational fluid dynamic method is applied to investigate the thermal and drag coefficient on the sharp nose cone with different cavity shapes. In order to simulate our model, the CFD method with SST turbulence model is applied to study the flow feature and temperature distribution in the vicinity of the nose body. The effect of depth and length of the cavity on the thermal characteristic of the nose cone is comprehensively investigated. In addition, the influence of the number of the cavity in the thermal performance of the main body is studied. According to our results, increasing the length of the cavity highly efficient for the reduction of the drag at Mach = 3. As the Mach number is increased to 3, the number of the cavity becomes a significant role and it is observed that case 9 with four cavities is more efficient. Obtained results also show that increasing the cavity depth declines the temperature on the main body. Our findings confirm that the main source of the expansion is the edge of the cavity.

This study investigates the impact of a shock generator on the fuel distribution of downward step fuel injection in a scramjet engine’s combustor. Computational simulations of CFD are employed to analyze the three-dimensional structure of multi-jets released in a downward-step configuration. The inlet is subjected to a free stream supersonic flow with $\mathrm{\text{Mach}}=4$, while hydrogen gas is injected at $\mathrm{\text{Mach}}=1$ through injectors situated on the vertical planes of the downward step. A shock generator with a $30^{\circ }$ angle is positioned at the top of the domain to induce a bow shock, which significantly affects fuel distribution. The research assesses the influence of free stream and jet pressure on the fuel distribution behind the jets. The results indicate that increasing the free stream flow strengthens the induced bow shock, leading to enhanced fuel jet interactions and improved fuel mixing behind the jets.

Acoustic analogues of black holes (dumb holes) are generated when a supersonic fluid flow entrains sound waves and forms a trapped region from which sound cannot escape. The surface of no return, the acoustic horizon, is qualitatively very similar to the event horizon of a general relativity black hole. In particular Hawking radiation (a thermal bath of phonons with temperature proportional to the "surface gravity") is expected to occur. In this note we consider quasi-one-dimensional supersonic flow of a Bose–Einstein condensate (BEC) in a Laval nozzle (converging-diverging nozzle), with a view to finding which experimental settings could magnify this effect and provide an observable signal. We discuss constraints and problems for our model and identify the issues that should be addressed in the near future in order to set up an experiment. In particular we identify an experimentally plausible configuration with a Hawking temperature of order 70 n K; to be contrasted with a condensation temperature of the order of 90 n K.

The lateral jet interaction on a slender body in supersonic flow was investigated by numerical simulation. The spatial and surface flow characteristics induced by jet interaction were shown. As a result, when the lateral jet is not in the longitudinal symmetry plane, the jet interaction causes asymmetric separation flow of surface and space, and destroys the pressure distributions of the slender body. With different angle of attack and circumferential positions of jet, the flow characteristic of the after body for jet in asymmetry plane changes greatly. The results with and without jet interaction also show that the far-field interaction played a major role in the lateral jet interaction.

To investigate the static pressure distribution characteristics of a flying-wing model, an advanced binary pressure sensitive paint (PSP) technique is introduced. It has low-temperature sensitivity and can compensate the errors induced by temperature. The pressure measurement test was performed in 0.6 m trisonic wind tunnel at angles of attack ranging from 0${}^{\circ }$ to 12${}^{\circ }$ in supersonic condition, adopting a low-aspect-ratio flying wing model. The binary PSP is sprayed on the upper surface of the model while pressure taps are installed on the upper surface of the right wing. Luminescent images of two probes are acquired with a color charge-coupled-device camera system and processed with calibration results. During the test, the surface pressure is measured by PSP and transducer, respectively. The results obtained show that the binary paint is of advantage to the surface pressure measurement and flow characteristic analysis. The high-resolution pressure spectra at different angle of attack clearly reveal the impact of leading edge vortex on the upper surface pressure distributions. The pressure measured by PSP also agrees well with the pressure tap results. The root mean square error of pressure coefficient is 0.01 at $\mathrm{\text{Ma}}=1.5$, $\alpha =0^{\circ }$.

Many projectiles tend to spin about their longitudinal axis while progressing in the forward direction. It helps in providing stability and a reference direction for guidance during their run. Many different projectiles employ a supersonic convergent-divergent nozzle to produce thrust for their forward motion; hence, the nozzle and overall whole propulsion system tend to spin about its axis of rotation. The main aim of this study is to observe the effect of spin on the nozzle. In this research, a converging bell-shaped diverging nozzle is numerically designed using a method of characteristics (MOC) for exit Mach number 3.21. Viscous simulations are performed for both two- and three-dimensional cases. The analysis is then performed with nozzle spinning about its axis of symmetry with a constant angular velocity of 10 revolutions per second. The analysis is repeated for the value of constant angular velocities to be 15 and 20 revolutions per second, and the behavior of flow with increasing angular velocity is examined. It has been observed that the exit Mach number and velocity decrease due to the radial protrusion of the boundary layer, and it has a negative impact on the performance of the nozzle. Moreover, the decrease of exit Mach number is in direct relation to increasing angular velocity.

The immersed boundary method (IBM) is an innovative approach for modeling flow with complex geometries and is more efficient than traditional method. In the present investigation, the shock wave propagation over one circular cylinder is simulated numerically with the Ghost-Body Immersed Boundary Method and high-resolution Roe scheme. To validate the IBM, a plane incident shock wave passing through a square cylinder is predicted and good agreement with previous experiments was obtained. Then based on our calculation, the reflection and diffraction processes of a shock wave passing through a circular cylinder were visualized and discussed in detail.

Velocity–temperature correlations in a high-temperature supersonic turbulent channel flows, including thermally perfect gas (TPG) and calorically perfect gas (CPG), are investigated based on the direct numerical simulation database [Chen et al., J. Turbul. 19 (2018) 365] to study the gas model effects. The results show that in fully developed turbulent channel flow, the Reynolds analogy factor remains close to 1.2 for both gas models. The “recovery enthalpy” is better than Walz’s equation to connect the mean stream-wise velocity with mean static temperature because it is independent with gas models. The modified strong Reynolds analogy for TPG is more accurate scaling than that for CPG, and the turbulent Prandtl number is insensitive to gas models. In addition, the influence of gas model on the probability density functions of stream-wise velocity and static temperature concentrate on the corresponding right tails.

In this study, the drag force and heat flux reduction mechanism induced by the aerodisk (with disks on its nose) with the freestream Mach number being 4.937 has been numerically investigated, and the simulations have been carried out by the three-dimensional Reynolds-averaged Navier–Stokes equations coupled with the SST $k-\omega$ turbulence model. The influence of the angle of attack on the drag and heat flux reduction has been analyzed comprehensively. The obtained results show that the drag force of the spiked blunt body can be reduced by the aerodisk, and the drag force decreases by 24.63%. The flow mechanism of the complex flow is drastically modified by the angle of attack, and this results in a strong flow asymmetry. This asymmetry becomes more and more obvious as the angle of attack increases. Both the pressure force and viscous force increase with the increase of the angle of attack. Moreover, both the lift and drag coefficients increase as the angle of attack increases, and the lift-to-drag ratio increases first and then decreases with the increase of the angle of attack. When the angle of attack is $6^{\circ }$, the maximum lift-to-drag ratio is close to 0.36.

In this paper, dynamic instability of functionally graded carbon nanotubes (CNTs)-reinforced composite joined conical-cylindrical shell in supersonic flow is analyzed numerically. The higher-order shear deformation theory is applied to describe the stress–strain state of thin-walled structure. The assumed-mode method is used to derive the finite degrees-of-freedom dynamical system, which describes the structure motions. The structure motions are expanded by using the eigenmodes, which are obtained by the Rayleigh–Ritz method. The trial functions, which satisfy the continuity conditions at the cylindrical-cone junction, are used to obtain the eigenmodes.

The properties of free vibrations of thin-walled structure are analyzed numerically. The dynamic instability of the joined conical-cylindrical shell in supersonic flow is analyzed using the characteristic exponents. As follows from the numerical study, the dynamic instability is arisen due to the Hopf bifurcation. The dependences of the supersonic flow critical pressure on the Mach number and the type of CNTs distribution are analyzed numerically.

In this paper, we establish the global existence and stability of a multidimensional conic shock wave for three-dimensional steady supersonic flow past an infinite cone. The flow is assumed to be hypersonic and described by a steady potential flow equation. Under an appropriate boundary condition on the curved cone, we show that a pointed shock attached at the vertex of the cone will exist globally in the whole space.

The aeroelastic flutter analysis of thick porous plates surrounded with piezoelectric layers in supersonic flow is studied. In order to aeroelastic analysis of the thick porous-cellular plate, Reddy’s third-order shear deformation plate theory and first-order piston theory are used. Furthermore, the plate is composed of two face piezoelectric layers and three functionally graded porous distributions core. Applying the extended Hamilton’s principle and Maxwell’s equation, the governing equations of motion are obtained. The partial differential governing equations are transformed into a set of ordinary differential equations by applying Galerkin’s approach. The effects of porosity coefficient, porosity distributions, piezoelectric layers, geometric dimensions, electrical and mechanical boundary conditions on the flutter aerodynamic pressure and natural frequencies of porous-cellular plates are investigated.

In this paper, flutter and divergence instabilities of functionally graded porous plate strip reinforced with graphene nanoplatelets in supersonic flow and subjected to an axial loading are studied. The graphene nanoplatelets are distributed in the matrix either uniformly or non-uniformly along the thickness direction. Four graphene nanoplatelets distribution patterns namely, Patterns A through D are considered. Based on the modified Halpin–Tsai micromechanics model and the rule of mixture, the effective material properties of functionally graded plate strip reinforced with graphene nanoplatelets are obtained. The aerodynamic pressure is considered in accordance with the quasi-steady supersonic piston theory. To transform the governing equations of motion to a general eigenvalue problem, the Galerkin method is employed. The flutter aerodynamic pressure and stability boundaries are determined by solving standard complex eigenvalue problem. The effects of graphene nanoplatelets distributions, graphene nanoplatelets weight fraction, geometry of graphene nanoplatelets, porosity coefficient and porosity distributions on the flutter and divergence instabilities of the system are studied. The results show that the plate strip with symmetric distribution pattern (stiffness in the surface areas) and GPLs pattern A predict the highest stable area. The flutter and divergence regions decrease as the porosity coefficient increases. Besides, the critical aerodynamic loads increase by adding a small amount of GPL to the matrix.

Supersonic flow offers a unique opportunity to follow the evolution of an aerosol on the microsecond timescale. The initial appearance of the droplets/particles and their further development can involve a number of nucleation steps: homogeneous nucleation from the supersaturated vapor to the (supercooled) liquid, nucleation of a solid phase from the supercooled liquid, and even heterogeneous nucleation of other species onto pre-existing particles. In this chapter, we present basic concepts from nucleation theory, the experimental techniques used to characterize the thermodynamic state of the flow as well as the size, state and composition of the aerosol, and the approaches used to determine nucleation rates. The nucleation experiments that have been conducted both within and at the exit of Laval nozzles are summarized, and the importance of experiments conducted in supersonic flows to advancing the field of nucleation is discussed.

The current work evolves a model for shear stress for a compressible mixing layer, with M_c being a parameter. This model is evolved with the data obtained from LES. It is seen that the shear stress to strain rate relation is remarkably linear in the self similar region of the mixing layer, albeit the proportionality constant depends on the prevailing convective Mach number. The model constants obtained from the simulations match with those which can be derived from experiments. By using this model, the growth rate obtained falls well within the range of experiments even at high M_c.