After the deflection of the wing control surface, flow separation is easily generated at the trailing edge of the wing, which will reduce the lift coefficient and the control surface efficiency. The rudder of the wing is aileron. If the lift generated by the wing is used to improve the efficiency of the control surface, the flow separation caused by the deflection of the control surface must be restrained. Using synthetic jet to change the flow state of boundary layer is the main method to solve the problem of flow separation. Synthetic jet actuator (SJA) has the advantages of no energy loss and simple structure. In this paper, a method of using synthetic jet actuator to suppress the flow separation at the rear of the wing when the aileron deflects is proposed, and the lift coefficient is obtained. The increase of aileron efficiency is calculated by the change of lift coefficient. The EPPLER555 wing with aileron deflection angle of 3°~9° is simulated, and the changes of lift coefficient and aileron efficiency under corresponding working conditions are obtained. The results show that the average lift coefficient of the wing is 0.5 when the deflection angle of the aileron is 3°~9° without SJA. After SJA employed, the lift coefficient will be greatly improved, and the control surface efficiency of EPPLER555 wing will be effectively improved, the lift coefficient will increase by about 20% to 0.6-0.7. For example, when the deflection angle of aileron is 4°, using a SJA with a maximum outlet velocity of 200m/s and an excitation frequency of 400/2π, the effective lift coefficient generated by the wing is 0.5931. Under the effect of SJA, the control surface efficiency of EPPLER555 wing will be effectively improved. The lift coefficient is reflected by the ratio of the change of lift coefficient after SJA employed to the lift coefficient without synthetic jet actuator.

The turbulent flow around a circular cylinder at the Reynolds number of 13100 is controlled using the combination of downstream splitter plate and upstream control cylinder as a new configuration. For this purpose, first a splitter plate with the length of L is attached to the downstream base of the cylinder and its length is varied to yield the optimum plate length in which the forces are minimum. The Reynolds averaged Navier Stokes equations is solved by a finite volume method together with a K-w SST turbulence model. Results show that the presence of the splitter plate has great impact on the flow characteristics and reduces the time average of the drag coefficient and amplitude of the lift coefficient oscillations. At L/D=1 the drag and lift coefficients of the system is reduced about 48% and 97% respectively. In the next part, a control cylinder with the smaller diameter than the main cylinder is placed at the upstream and the center to center ratio is varied. Results show that the control cylinder reduces the system drag coefficient in all the investigated distances while reduces the lift coefficient just in some bigger distances. At C/D=4 the drag and lift coefficients of the system is reduced about 35% and 11% respectively. In the last part, the simultaneous combination of upstream control cylinder and downstream splitter plate in their optimum configurations is used. Results show that this configuration significantly reduces the drag and lift coefficients of the system compared to the main cylinder. Furthermore, the performance of the new configuration in reducing the drag force is better than those of each of the splitter plate and control cylinder alone. At the condition with C/D=4 and L/D=1 the drag and lift coefficients of the system are reduced about 57% and 63% respectively compared with those of the main cylinder.

In this paper, the effect of the Gurney flap parameters such as the angle of attack, Reynolds number, angle and height of the flap and its location from the escape edge on the lift coefficient of a symmetric airfoil is considered with the help of simulation in computational fluid dynamic software of Fluent. The turbulence model k-ε is used for the two-dimensional domain. Also, the value of the lift coefficient is introduced as a function of effective parameters by the design of experiment (DOE) method and using the backward elimination regression model which is a statistical method for selecting the model and estimated error terms. The value of the airfoil lift coefficient can be determined and predicted by the obtained function. The numerical values derived from the function of the lift coefficient resulted from the design of experiment method are in good agreement with other valid papers. The results show that at the constant attack angle by increasing Gurney flap height, lift coefficient increase. On the other hand, at the constant height of the Gurney flap, this coefficient decreases with increasing angle of attack. Moreover, the lift coefficient increased by increasing the distance of the Gurney flap from the airfoil escape edge at a 90-degree angle and 1.5%, as well as increasing the Reynolds number at a constant height of a Gurney flap.

With the advent of morphing airfoils, the aerodynamics of wind turbines and wings underwent many changes. In this study, the aerodynamic coefficients of morphing airfoil based on NACA 0015 are optimized in the range of Reynolds number 105 to 106 and the angle of attack 0 to 12 degrees using Artificial Neural Network (ANN) and Genetic Algorithm(GA). First, the airfoils were created in MATLAB software by random control points and mesh generated in Gambit software, then in two-dimensional Ansys software were simulated. The simulation results, including lift and drag coefficients, separation point and pressure center, with control points were used to train the Artificial Neural Network (ANN). The trained function is given as an input function to the Genetic Algorithm(GA) to optimize the desired coefficients.Lift coefficient, the center of pressure, separation point and lift to drag ratio were optimized as a single objective, In single-objective optimization, the lift coefficient was increased by 18% using the morphing airfoil. Also, the lift coefficient and the center of pressure, the lift coefficient and the drag coefficient were optimized as the dual-objectives optimization. In the optimization of the dual objectives, lift and drag coefficients were controlled by 0.8 and 0.03, respectively, by the morphing airfoils.

Numerical simulations of the airflow around a hatchback and a sedan vehicle without and with spoilers are carried out, besides, its effect on drag and lift coefficients are investigated with and without crosswinds. The effects of crosswind on aerodynamic forces are considered and its results are compared with the case without considering the effects of crosswind. For this purpose, the steady-state three-dimensional Navier-Stokes equations are solved by the Simple Method. Moreover, for turbulence modeling, the Realizable k-e model is implemented. The spoiler angle and its length are changed for both car models; furthermore, the effects of two spoilers on drag and lift coefficients are investigated in detail. All cases are simulated with and without crosswind. The results show that the impact of the spoiler for without crosswind conditions to decrease the lift coefficient in both models is significant; in addition, the drag coefficients are reduced for some cares. It can be concluded that the increase of spoiler length for both sedan and hatchback vehicles can increase the downward force and vehicle stability.

In this paper, several models of structure-fluid coupling are employed to investigate on the vibration behavior of the structure. Using basic equations of vibration and employing a mathematical model, a single degree of freedom system is analyzed. Some parameters of the mathematical model are obtained from test. To examine structure-flow interaction, coupled system of nonlinear second-order differential equations, including vibration equation of structure and van der Pol's equation, are solved synchronously. Lift coefficient is obtained by solving the coupled equations. Numerical solution is accomplished for three coupling models: displacement coupling, velocity coupling and acceleration coupling in each of which force is function of displacement, velocity and acceleration respectively. Time response, lift coefficient and vibration amplitude in steady state are obtained and plotted. Phase angle between the structure motion and lift coefficient change considerably when passing the locking zone which is well coincident with experimental results. Steady state vibration amplitudes for mentioned models are verified by comparing with experimental results.

Determining characteristics and behavior of fluid flow and obtaining the aerodynamic coefficients are important and considerable tasks in investigation of flow over the bodies. In this research, leading edge radius effects on the flow behavior over a delta wing with a sweep angle of 50 degree at different Reynolds numbers between 0.5×106 and 2.5×106 are investigated numerically. Three different leading edge models including one sharp and two round (having different radius) are studied. Validation of the results is performed using available experimental results which showed good agreements. Leading edge radius effects on the aerodynamic coefficients and size and location of the vortices are investigated. Based on the obtained results, on a fixed angle of attack, increasing the wing leading edge radius reduces the moment coefficient. Also, it decreases size and strength of the preliminary vortex and magnitude of the lift and drag forces. When the Reynolds number varies Between 0.5×106 and 0.97×106, the lift coefficient tends to experience a 6 to 8 percent raise, while at higher Reynolds numbers, no considerable raisings observes. Also, at angle of attacks of 20 and 25 degrees, by increasing the Reynolds number, the lift coefficient graph experiences a noticeable reduction. At the sharp leading edge delta wing, stronger vortices generate which results higher lift coefficient. By rounding the leading edge, due to weakening the aforementioned vortices, the lift coefficient decreases.

In this numerical investigation, the effect of upstream and downstream interference of two identical tall buildings was evaluated using the computational wind engineering (CWE) method. In order to simulate three-dimensional turbulent wind flow with Reynolds numbers in the range of 1.4×〖10〗^4<Re<8×〖10〗^4, the Large Eddy Simulation (LES) model was used. Mean and fluctuating coefficients of drag and lift and pressure distribution were used as the main criteria to evaluate the aerodynamic response of the principal building in different conditions of interference. Streamlines and vorticity contours were presented and their relationship to aerodynamic results was interpreted to provide a better understanding of the physics of the problem. According to the results, the shielding effect of the interfering building in most interference cases led to a reduction of the mean drag coefficient of the principal building compared to the isolated case. Compared to the isolated building, depending on the location of the interfering building, the fluctuating lift coefficient either increases or decreases.Upstream interfering buildings with sufficient distance to the principal building lead to an increase in the fluctuating lift coefficient of approximately about 50 percent. While, the fluctuating lift coefficient of the principal building decreases up to 37 percent, due to the closely spaced tandem interference states. While the mean pressure coefficient at the windward surface is not significantly sensitive to the interference states, it is strongly influenced by the different states of the interference at the lateral and leeward surfaces. Aerodynamic parameters were less sensitive to Reynolds number variations due to the fixed position of the flow separation in the sharp corners of the rectangular section and the zero angle of attack in this study. Changes in the Reynolds number resulted in variations of about 4 to 10% in the drag and lift coefficients of the principal building.

When a flying vehicle is approaching a watery or earthy surface, the flow pattern around it is changed that is called the ground or surface effect. In this study, the phenomenon of ground effect and its effects on aerodynamic coefficients and flow pattern around NACA0012 and LH37 airfoils are numerically investigated. The analysis is done for statically and dynamically airfoils with plunging motion at subsonic incompressible flow regime. The Navier-Stocks governing equations are used with k-𝜔 SST turbulence model. At first the effects of ground effect on lift coefficient of airfoils are studied in various distance from surface, statically. Then at each position of airfoils from the surface the lift coefficient behavior of airfoils at sinusoidal plunging motion with the specified amplitude and frequency is investigated. the statically results show that the lift coefficient of airfoils and pressure distribution over them are changed when they approach the surface with respect to far from it, which is seen as decreasing to a certain height and then increasing it. Dynamically analyzes also indicate a change in the oscillation amplitude of the lift coefficient and the existence of a phase difference at the points of achievement of minimum and maximum lift, when the airfoils are an approach to the surface. The streamlines also showed the changes in flow field patterns around the airfoils, when they approach the surface.

In this paper, the Semi-Empirical and numerical methods that can be used to investigate the effects of dynamic stall are compared with each other, and the capabilities of the methods are studied. The experimental measurements have been used in order to compare the methods. The Semi-Empirical Leishman-Beddoes (L-B), Snel and ONERA methods have been used, and the finite volume method was being used for numerical simulations. The lift coefficient was being calculated by all the methods at various conditions, and the drag coefficient had been computed by the numerical and Leishman-Beddoes methods. The parameters that have been used in order to compare the methods, are the maximum lift coefficient value, the angle of attack of the largest lift coefficient, the error at upstroke phase and the error at down stroke phase. The results show among the semi-empirical models; the L-B method has the highest precision to predict the lift coefficient, and although the numerical method can investigate the flow with more details, but the error percentage at the down stroke phase is higher than expectations. The results from the drag coefficient modeling show that the numerical method can predict this coefficient better than the L-B method. The results also can help other researchers to select the best dynamic stall model in order to investigate the wind-turbine aerodynamics.