عباس طربی

Lattice fin is an aerodynamic control surface with an outer frame and square or diamond-shaped grid structure of secant septum. The purpose of grid fins is to provide a level stabilizer or control level of the missile while reducing the hinge moment at speeds in which many forces enter the fins. In this research, with computational fluid dynamics method, the septum and density effect of lattice fins on the aerodynamic coefficients of the missile in steady state and supersonic flow regime has been studied. Simulation of lattice fins in this paper, at first have two types of 3D sharp diamond and flat plates, then compared with the near experiment results that were available to them and after that selecting the 3D sharp diamond septum. At last variant of missile aerodynamic coefficients with 50% increase and decrease of septum density fin (unchanged in the overall dimensions of the fin frame) was compared to the reference state. Diagram of axial and vertical force coefficients, as well as the pitch moment and center of pressure in the assumed missile, are given for various angle attacks in Mach 3. Numerical flow simulation results around two types of lattice fins indicate the accuracy of numerical solution in calculating the flow complexities on lattice-shaped geometries and also matching of 3D sharp diamond septum results with the experimental results of wind tunnel. Also, calculations show that by increasing and decreasing the 50% density of septum in a fixed frame, the position of center of pressure can be changed by about one diameter to the front or rear of missile.

In this article, investigating fin type geometry effects with numerical analysis on aerodynamics coefficients in steady state and supersonic flow. In fact the Grid Fin is an aerodynamic control surface with a grid structure (square or rhomboid) of thin septum that has many advantages and provide high levels of stabilizer or control surface for the missile. The simulated lattice fins in this paper have rhomboid partitions and investigating the variance of aerodynamic coefficients of the missile with an increase of 25% & 50% in the span, an increase of 50% & 100% in chords and an increase of 1.5 times the width with the constant fin thickness assumption (no change in general dimensions of the fin frame). Initially, the axial and vertical force coefficients as well as the pitch moment and center of pressure on missile were calculated at various angles of flight on Mach 3 in main grid fin and compared with numerical and experimental results in reference articles. The results of the comparison show the numerical resolution accuracy on calculating the flow complexities in the missile latticework with the experimental results of the wind tunnel. Further, by increasing the span, chord and lattice width, the aerodynamic coefficients variant of the missile were compared to the reference missile. Calculations show that by increasing the span and chord, all the aerodynamic coefficients will increase to the extent that they will be mentioned, and with increasing width with constant fin thickness, Drag coefficient will be reduced to a small extent and the remaining aerodynamic coefficients will increase.

In this research, procedure of aerodynamics design for hypersonic wind tunnel test section is investigated by using computational fluid dynamics. The type of test section is chosen closed-type free jet. First, conceptual design is done by using the available statistical data in the hypersonic wind tunnel test section in the world. Then, quality of design is investigated by numerical simulation by solving Navier-Stokes equations in compressible, viscous (k-w SST model) and axisymmetric form. With changing the parameters of the test section, the test section dimensions have been optimized. The important parameters in test section are length of the test section, the test section diameter, the diameter of the inlet diffuser, the diffuser and the nozzle indentation inside the test section. Design criteria are chosen the formation of uniform flow, better swallow the slip line and the flow deflection in the test section. Finally, dimensions, specifications and aerodynamics functional requirements of the hypersonic wind tunnel test section have been determined at Mach 5 to 7. The effect of cone and curved nozzles is also investigated for flow quality in test section of wind tunnel.

Reduction of fossil resources, significant growth of oil and natural gas price and greenhouse gases release enhancement have caused more usage of renewable energies as an unlimited, free, clean and environmentally friendly resources. Due to the vast expanses of wind farms in Iran, there is a good basis for using wind energy as an accessible, unlimited and environmentally friendly source. The variable wind speed and the randomness of the nature of the power generated by wind turbines have made it possible for energy experts to provide sustained and permanent load with wind energy hybrid with other renewable sources. In this regard, since 2007, many studies have been done in Iran on modeling and optimizing hybrid power generation systems based on wind turbines. HOMER software is considered as one of the main tools for modeling these systems by Iranian designers. The present study reviews and analysis these indigenous studies. Reviews show that more than half of the research has been published on how to provide power through the use of the hybrid wind, PV, and generator systems.

In the present study, the numerical solution of the Ansys Fluent software has been used to calculate the sound produced by the high-speed flow on a cylinder using the Lighthill acoustic analogy. The calculations were carried out on a cylinder (part of the landing gear) at a speed of 70 m/s (take-off and landing speeds of airliners). The problem is initially caried out as a regular unsteady numerical solution. During the solution, aerodynamic noise data sources are stored as inputs of acoustic analyzes in files. Then, by solving the acoustic equations, the volume of produced sound (in decibel) is calculated at points that are pre-defined as the microphone in the desired coordinates. The purpose of this study is to study the ability of Fluent solution to calculate the sound generated by the flow, in addition of using a method for estimating the amount of sound increase by increasing the length of the cylinder. In the other words, due to the timing of the numerical solution, one can calculate sound generated by small length cylinder, and then, using engineering approximation, it estimates the sound of the flow around the larger-length cylinder. After the necessary calculations, results are provided as sound pressure level curves using the acoustic analogy and fourier spectral analysis. The results show that large eddy simulation turbulence model is most appropriate model for acoustic simulations. Also, the approximate method for evaluating the effect of increasing the length of the cylinder is in good agreement with the experimental results.

The ability to control the flow, is one of the basic needs of Fluid Mechanics that constantly pursued by researchers. One of the new methods in this area, is using Dielectric barrier discharge (DBD) plasma actuators that by injecting momentum into the boundary layer, causing a delay in the phenomenon separation. The main object in this work was to help to optimize the electrical parameters to obtain stranger vortex and more effective ionic wind created by steady and unsteady plasma actuators on the air through the flat plate. For this reason, simulation is done for a flat plate with the compressible 5 m/s velocity airflow. The time averaged velocity profiles of the ionic wind show that averaged velocity come more and the position of the maximum velocity come near the surface by increasing the excitation voltage and frequency. The power, of the vortices that are shed form the unsteady actuator, increases by increasing duty cycle percentage. Our results on the ionic wind velocity on different position on the flat plate indicate that the maximum averaged velocity occurs in downstream of plasma actuator.

In this paper, FLUENT software capabilities are used to calculate and compare the sound produced by the flow of air through several geometries with different cross-sections. The method used in this simulation is the combination of the large eddy simulation turbulence model and Ffowcs-Williams and Hawkings acoustic model. Three cross-sections of circle, square and triangle are studied. The aim of this study, in addition to assess the ability and accuracy of numerical calculation of the farfield sound, is to find the maximum sound pressure level of each simple cross section. Two microphones are defined over long distances of the numerical domain to compare the numerical sound calculations with experimental results. After the necessary calculations, results are provided as sound pressure level curves of turbulence solver using the acoustic analogy at microphones locations. Respectively, the received sound pressure at the microphone farther- located is in lower quantity and the arrival time of the sound at that location is longer. The results show that the maximum sound pressure level, at the same velocity, belongs to circular, triangular and square cross-sections respectively.

In an attempt to achieve efficient operation with respect to design and combined thermal and electrochemical perspective, a computational model of tubular SOFC is presented. The FLUENT software was used for the development of the model. The particular model, after validation against experimental observations for selected benchmark cases, was demonstrated to be a basic model for further research in this area. For validation of obtained result, i.e. COMSOL Multiphysics, was used which is based on Finite Element Method (FEM). The results showed that the difference of them is less than of 6% for power density and less than 7% for output voltage and so both software are acceptable.

Lateral jet control systems are being considered as attractive alternatives to conventional control systems in recent years. In present study which is divided in two parts, the effects of lateral jet interaction with supersonic cross flow on aerodynamic behavior of a standard projectile at zero angle of attack has been studied. In the first part, results of the effects of parameters such as jet location, Mach number and nozzle type on pressure coefficient, drag coefficient, drag force and pressure distribution on the fins is presented and analyzed. In the second part, longitudinal static and dynamic stability coefficients of the projectile in presence of lateral jet has been achieved and evaluated according to the mentioned parameters. According to the results, jet location is the most effective parameter. In the first part, the pressure distribution on the fins is much dependent on jet location. Effect of Mach number on pressure coefficient, drag force and drag coefficient is also significant. Besides change of the pressure distribution on the fins comes more into sight at the final locations by variation of Mach number. In the second part, lateral jet effect leads to decreasing longitudinal static stability. Increasing the Mach number is also results in decreasing longitudinal dynamic stability and jet displacement make nonlinear behavior over pitch damping moment coefficient, therefore choosing proper jet location is depend on desired parameters of designer. According to the results, effect of nozzle type has been insignificant for all cases.

Boundary layer flow compressibility, and permanently with sticky mass transfer (suction or injection flow) on an axial symmetric Bdnhhay study and numerical study is. Newton fluid gas (air) and considered a constant suction or injection flow rate as national or local by creating holes in different parts of the body surface is applied. General equations Navyr - Stokes equations with the approximate boundary layer, supersonic flow conversion and using Ghyrtshabhy Mhasbh distribution and integral to the heat transfer surface and nose profile to determine if the numerical boundary layer are analyzed. Governing equations with boundary conditions using the conversion Prabstn - Elliot coordinate the flow of lines and then transferred to the system and Exposure Nonlinear Equations with Parabolic the inclination toward nature as step by step solution Grdydhand To determine the flow profile, three-mode Dyvarh rigid, porous Dyvarh with suction or injection flow is considered. Numerical Flow Simulation on a promontory with a vertex angle of 20 degrees, Mach numbers in the 5 / 2, 5 / 3 and 5 / 4, and 35 kilo Pascal pressure 30, and 293 and 311 degree Kelvin temperatures has been done. In order to verify numerical simulation results with the numerical values of others and also the results of implementing software that Flvynt being compared are completely satisfactory.

In this paper, the implementation of the variable-span Morphing wing(VSMW) on the wings of a UAV to achiev the optimal performance for various flight regimes is studied in wind tunnels. For the implementation of this plan, a morphing wing with its related mechanism is designed, built and tested. Tests were carried out at a low- speed wind tunnel with speeds of 35, 60 and 80 m/s in the range of Reynolds Number between 250×〖10〗^3 to 620×〖10〗^3. For the implementation of the morphing wing plan, the maximum deviation of wing span relative to Basewing (BW) is 38 percent and change of aspect ratio is 74% percent. Experimental results showed the improved aerodynamic properties of VSMW compared to BW. These improvements are in the form of reduction of induced drag, thrust required and increasing of lift coefficient, aerodynamic efficiency, endurance and rang. According to the results, endurance and range for VSMW comparison to BW, respectively 17 and 5 percent increased. The experimental studies show that the maximum aerodynamic efficiency of the wing model occurs at 4 degree angle of attack with the maximum value being 14.8 for the base wing, 16.3 for the morphing wing with 50 percent (5 cm) extension in wing span, and 17.4 % for the morphing wing with 100 percent (10 cm) span extension. This research also shows how VSMW can be used as an alternative method of roll control instead of the regular method (i.e. using aileron control surfaces) as it provides more roll control power than the regular method.

In this research, experiment and computational fluid dynamics (CFD) are used to assess the performance of UAV with variable-span and sweep morphing wing under low speed conditions. Both wing area and aspect ratio are changed due to variations in span and sweep, whereas structure of the variable-span and sweep morphing wing remains constant. In this study, the numerical results of Fluent software and experimental data are presented. Results are achieved under a low wind speed (50, 60 and 70 m/s). In this case, full extension represents a 30% (10 cm) increase in wing span and 36% (12 deg.) in sweep angle relative to the original wing, with no extension. The results of this study show that the morphing wing is capable to improved aerodynamic efficiency, increased both range and endurance, reduced induced drag and in general reduced thrust required. According to experimental and numerical results, the use of morphing wing can increase the range by 13.6% and 13.5%, also, endurance of the vehicle by approximately 8.855 and 8.17%, respectively. The results of this study show that the maximum value of lift-to-drag ratio occurred at 6 degrees angle of attack and a speed of 70m/s. These results demonstrate that the use of morphing wing improve the lift-to-drag ratio by 10% compared to original wing. Finally, the numerical simulations are compared and show good agreement with the experimental results. This research also showed how morphing concept can be used as an alternative method for roll control.

The purpose of this investigation is consideration of aerothermodynamic phenomena over blunt and slender pointed bodies in hypersonic flows. Such results may be obtained by numerically solving the full Navier-Stokes equations or one of their various subsets، such as PNS equations. However، due to the excessive computer storage and CPU requirements، most people use their simple forms. In this work، a computer code was developed that uses an implicit finite difference method to solve non-similar، full compressible boundary layer equations for laminar، transitional، and turbulent flows over blunt and slender pointed bodies in hypersonic flows. Different test cases were studied and the related results were compared to benchmark CFD and experimental data showing relatively close agreements.