مجله مکانیک سیالات و آیرودینامیک
سال دهم شماره 2 (پاییز و زمستان 1400)

Cargo parachutes are commonly used to deliver heavy instruments to areas with difficult access. Perceived, good stability and low descent rate are two essential parameters for a recovery system. According to the above factors, one of the popular cargo chute configurations is a combination of a large chute as the main chute and a smaller one as a stabilizer chute. Since there is few scientific research that has studied on aerodynamic characteristics of this cargo chute configuration, a three-dimensional numerical simulation was performed here to investigate this kind of parachute system. To this end, main chute dimensions were computed based on mission definition (weight and descent rate). In the following, the effects of cargo distance from parachutes are studied for an optimal design of the main and stabilizing parachute. At five intervals of 1, 2, 4, 6, and 8 times the diameter of the main parachute, between the cargo and the main parachute distance, the values of drag coefficients and lateral forces for the main and stabilizing parachute are reported.

In this study, the effect of roughness and stiffness on the aeroelasticity of an oscillating airfoil during turbulent unsteady transonic flow has been studied. In this simulation, the finite volume method is used to discretize the equations to solve the Navier-Stokes equations. In this pressure-based algorithm, a high-resolution scheme for convection term and 𝜿-ε turbulence model are used. For computing convection terms, a Normalized Variable Diagram technique is used. Here the technique of inlet velocity vector oscillation is applied. In addition, a modified 𝜿-ε model for compressible flow is applied to simulate Navier Stokes equations. The two-dimensional motion equations are obtained from the Lagrangian equations, which are combined with the aerodynamic equations. The results of validation show that the extracted data has a desirable accuracy. Furthermore, the FSI results show that, for rough airfoils, the strength of the shock wave is weakened, the shock wave moves to the trailing edge, and the oscillation of the airfoil is reduced. Also, with increasing structural stiffness, the damping of oscillations increases, and drag decreases.

In this study, by considering jacobian matrix based on conservative variables, eigenvalues, eigenvectors, and characteristic values for three types of preconditioners which are introduced by Turkel, Choi&Merkel and Eriksson are drawn in a unified mathematical manner. For this aim, these preconditioning methods are implemented in two-dimensional density-base “Roe” upwind scheme on unstructured meshes for Euler equations. Accuracy and rate of convergence are examined by external computing flow over NACA0012 airfoil, three-element 30P-30N airfoil, and internal flow over the bump for different flow conditions. This study shows that the application of preconditioning schemes not only increases the rate of convergence for compressible and incompressible flows dramatically; but also improves accuracy for incompressible flow in comparison with the classical method. This study also indicates that all preconditioning schemes provide approximately the same accuracy, but in terms of convergence rate, the Turkel preconditioning scheme provides a better rate of convergence among all the aforementioned preconditioned matrixes.

In present study, the effect of vortex generators on the aerodynamic forces of a helicopter's main rotor in hover is investigated. Vortices around blades have a significant effect on noise and aerodynamic forces. Using vortex generators is an appropriate method for decreasing the influence of flow separation and vortex around a rotor. In this study, vortex generators include four different arrangements. The numerical simulation procedure of main rotors of the helicopter blades is done by using Fluent software. The mesh grid is used in the form of unstructured. A validation with Cardona and Tang results is done. For this reason, NACA 0012 airfoil and an attack angle of 8° are used for the blades of the main rotor. It can be revealed from the results that using vortex generators, a decrease in vortex power and dimensions of transverse vortices has occurred. Also, the results show that the thrust coefficient and torque coefficient of the helicopter blades compared to the non-vortex generators mode has increased and decreased, respectively.

In this paper, CFD aerodynamic simulation of turbulent compressible fluid flow around high lifting control surfaces of a wide body heavy commercial aircraft in different flight phases using Spalart-Allmaras single equation density based to extract of the lift, drag and mean pressure aerodynamic coefficients has been carried out. Modelling of the wing of Airbus A380 in the actual dimensions based on super critical airfoil SC(2)-0610 using 3D design software SolidWorks to generate 2D and 3D geometries has been done. The unstructured mesh in 2D and 3D according to the different configurations of 2D airfoil/3D wing control surfaces has been done using ANSYS Workbench meshing tools and then the CFD modelling for a turbulent compressible subsonic air flow regime in 2D and 3D using ANSYS Fluent is done. Mesh independence study, model validation and comparison of the results for the 3D wing is done. The effect of changes of different configuration of slat and flap lifting devices in the leading and trailing edges and for different angels of attack before stall of the wing according to different flight phases (take off, cruise and landing phases) on the aerodynamic coefficients in the 2D/3D turbulent flow regime have been investigated. Contours for pressure, velocity, Mach number distribution and velocity vectors around the A380 airfoil and wing control surfaces has been presented.

Accurate aerodynamic coefficients and stability derivatives estimation is one of the feature of aeroballistic tests. This can be done by measuring the linear and angular coordinates, velocities and other relevant parameters from different sensors. The important question is the accuracy of the results, the method correctness and the coefficients estimation algorithm. Therefore, the effect of any measured parameters error on the estimated aerodynamic coefficients should be considered and analyzed.In addition, in order to investigate the effect of errors on the estimation of aerodynamic coefficients, a set of aeroballistic tests for measuring the sphere drag coefficient with error of less than 3% in this study. This shows the capability of high precision coefficient measurements in such tests. The errors involved in aeroballistic tests are mainly due to sensors, model manufacturing, data collection and the estimation algorithm. By modelling the error origins, the final error of the system can be predicted and analyzed. This investigation shows effect of each sensor error on the modelled collected data plus other flight simulation data that would help in designing the actual experiments.

Flow electrochemical batteries such as Al-Ago Batteries are high-capacity batteries that the electrolyte flows between the anode and cathode plates and the battery cell is in a close system. Due to the heaviness of such batteries, efforts are done to reduce their volume and weight so the distance between the anode and cathode is minimized. On the other hand, being to close to the anode and cathode increase the risk of short-circuit the battery. So, separators between the anode and cathode were used to prevent them from touch each other. Although preventing short-circuit, the separators are located in the direction of the electrolyte flow, preventing the desired flow between the anode and cathode plates. The flow wake due to their obstruction on the flow path effects the active area of the reaction and thereby reduces battery efficiency. On the other hand, the cross-section shape of these separators has a great influence on the hydrodynamics of the electrolyte, wake length and the active area of the reaction. In the present study, the electrolyte is considered as single-phase, steady and the effect of the number and arrangement of the separators are studied by four shapes of circle, square, Lozenge and triangle cross-sections on the electrolyte flow and the active area of the reaction.

In the present study, the entropy generated due to the conjugate heat transfer of the hybrid nanofluid inside the K-shaped chamber under magnetic field and uniform heat absorption/generation is investigated. The simulation was performed by writing computer code in Fortran language using the lattice Boltzmann method. Variations in Rayleigh number, volumetric fraction of nanoparticles, Hartmann number, heat absorption/generation coefficient, thermal conductivity ratio, chamber aspect ratio and type of magnetic field applied have been evaluated as the main variables of this study. The findings showed that the flow strength, heat transfer rate and entropy produced could be reduced by applying a magnetic field. A lower reduction of the average Nusselt number is achieved by non-uniform application of a magnetic field. Increasing the heat absorption/generation coefficient due to increasing the set temperature leads to decreasing the mean Nusselt number, which this influence increases with increasing the Hartmann number. Addition of nanoparticles to the base fluid in which the conduction of the phenomenon is predominant, increases the rate of heat transfer. Heat transfer is a function of the ratio of thermal conductivity and Rayleigh number that increasing these two parameters increases the convection effects, and in this case, the effect of increasing the Hartmann number is more pronounced. Increasing the chamber aspect ratio leads to a decline in the mean Nusselt number and entropy production, but the effect of adding nanoparticles is greater in this case. Entropy production decreases with increasing Hartmann number and increases with Rayleigh number and heat absorption/generation coefficient.

rapid population growth will increase the need for renewable energy resources. On the other hand, the extent of pollution from fossil fuels has made life on Earth difficult. However, the need to choose a suitable, cheap and clean alternative to fossil fuels is obvious. One of the proposed energy sources is electrical energy generated by fuel cells, which are currently a suitable solution due to high efficiency, non-pollution of the environment and the use of hydrogen as fuel. In this research, a solid oxide fuel cell with two different geometries is simulated in three dimensions. The equations governing the performance of the fuel cell, including electrochemical, momentum, mass transfer, and energy, are coupled using a finite element code defined, solved, and investigated. The results showed that the tubular geometry with the same dimensions and mechanical characteristics has a better performance than the planar type. It was shown that solving the energy equation and non-uniform temperature distribution reduces the power density of the fuel cell by about 7%. It was also found that cathodic pressure changes have a greater effect on fuel cell performance than anodic pressure changes. In the end, the results showed that increasing the thickness of the anode has a significant effect on increasing its performance compared to increasing the thickness of other fuel cell components.

In this paper, the effect of variation in the expansion ratio of the convergent-divergent nozzle on the performance parameters such as specific impulse, nozzle output velocity and output temperature is investigated using thermodynamic relations for different propellant. Then, three nozzles with different expansion ratios are manufactured and their thrust force with three different propellants is measured using experimental tests. The results show that with increasing the area expansion ratio, specific impulse, nozzle output velocity, output Mach number and thrust coefficient increase nonlinearly and the nozzle output temperature decreases. In addition, it is observed that with increasing the specific heat ratio of propellants, the output Mach number increases and the thrust coefficient and output temperature decrease. Also, with increasing the specific heat ratio and increasing the specific constant of the gases, the specific impulse and the nozzle output velocity increase. Furthermore, the thrust force increases with increasing nozzle expansion ratio and decreases with increasing propellant heat ratio. Finally, by comparing the thrust force obtained from the thermodynamic relations and their counterpart measured thrust force, the accuracy of the calculations is confirmed

In this investigation, grid generation difficulties in CFD analysis of internal combustion engines (ICEs) have been studied. Grid generation is primitive, extremely complicated and time consuming task in ICEs simulations. The KIVA open source code is one of the most popular CFD solvers in ICEs simulations. Grid generation for this solver is the most complex and difficult task. In the present work a methodology has been developed for rapid grid preparation in KIVA-3V code by using commercial grid generator, ANSYS ICEM CFD. In this methodology, all geometry details of ICEs, including valves and intake/exhaust ports, are taken into account. In this paper, the focus is not on modifying the KIVA code original mesh generator, however, the most popular grid generation and dynamic mesh management errors and some bugs in ICEM CFD and KIVA have been inspected and explained. By using the procedure described here, many of grid generation difficulties will be eliminated and grid generation time will be greatly reduced. It should be noted that many of features in this paper are general and can be used for geometries other than ICEs.

In this study, the growth of glaze and rime ice along the UAV’s wing span was studied, the cause of the formation of these ices on the surface and the effect of ice accretion on the aerodynamic performance of the wing was investigated. For this reason, a rectangular wing with NACA 0012 airfoil section at an angle of attack of 4 degree was studied at two different temperatures. A pressure-based solver and the Spalart-Allmaras turbulence model were used in commercial software. Calculations were performed at Re = 3×106. The results of the ice growth pattern indicate that there was no difference between the ice thickness on the wing span from root to middle, but from the middle to the tip, due to the increase in flow velocity, the rate of collision and the accumulation of droplets in the area increased. Also under glaze ice conditions, ice forms near the trailing edge due to the growth of the boundary layer. This calculation also proved the accuracy of this claim by performing similar calculation in the inviscid condition and the lack of ice growth near the trailing edge. On the other hand, the vortex phenomenon that occurs on the tip of the three-dimensional wings causes part of the droplets to hit the tip of the wing, resulting in the growth of a small amount of ice in this area. study of lift and drag coefficients showed that ice formation reduces the aerodynamic performance of the wing.