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In this paper a continuous-time state-space aerodynamic model has been developed based on the boundary element method. First, boundary integral equations for unsteady potential subsonic flow around lifting bodies are presented with emphasis on a modified formulation for thin wings. Next, the BEM discretized problem of unsteady flow around an arbitrary wing is recast in the form of a state-space model using some auxiliary assumptions. To validate the proposed model, its predictions for unsteady aerodynamic coefficients due to various unsteady flows about different wing geometries were compared to the verified results of the direct boundary element solution and good agreement was observed. Because of the resulting aerodynamic model has been constructed in the continuous-time domain, it is particularly useful for optimization and nonlinear analysis purposes. Moreover, its state-space representation is the appropriate form for an aerodynamic model in control applications.

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This work is the result of a research on the lifting forces during upward bird flight via modeling and manufacturing dynamical structures resembling bird wings of sizes between half to about 2 meters. The variables in this work included the wings sizes and their oscillation frequencies. In our formulations the lifting force and the consumed power at the beginning of a bird flight in a fixed frequency is proportional to the fourth and fifth power of the wings sizes, and for fixes sizes is proportional to the second and third power of the frequency, respectively. The lift force here is taken to be of two forms. The first is the very form relevant to the manufactured and used wing systems in the present work. In the second form the wings are assumed to stay horizontal during their vertical periodic motion. The extent of validity of these formulations when practicing for our manufactured wings, and for the real functioning of bird wings as well, has been the most important question in the present research. As far as the lifting force is concerned, the extrapolation of final results seems to be in consistence with the sizes relevant to human “bird-like” flight. However, provision of the needed power necessitates requirements to be thought of deliberately for restoring the energy in an effective way.

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In the recent years, wind energy had a faster growth compared with the other renewable energies. The interaction between fluid and structure becomes more important as the wind turbine size and its power production capacity increases. In the present research, the effect of wind speed and blade materials on static deformation of a small size horizontal axis wind turbine blade has investigated. The shaft torque and root flap bending moment values obtained from simulation are in a good agreement with experimental data. Results demonstrated that the deformation of the blade increases as the wind speed grows although the increase rate has declined in the mean wind speed range because of the occurrence of separation phenomenon on the blade surface. The effect of blade components materials on blade deformation was investigated and the least deformed configurations were introduced. The thickness of the designated blade components has been investigated by means of the maximum strain theory. The final thickness of the skin, spur and root was estimated by 2.1 mm, 2.8 mm and 10 mm respectively which are 30% less than the primary one.

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The aerodynamic characteristics of nine configurations of supersonic continuous deflectable nose guided missiles have been investigated. Then the optimized geometry is achieved based on the maneuverability from aerodynamic and flight dynamic point of view. The studied configurations consists of a spherical nose tip, a tangent ogive, one set of stabilizing tail fins and a cylindrical body that its mid-section is flexible to form an arc of a circle. So the cylindrical body consists of a fix part in vicinity of nose, middle flexible part and main body with stabilizers. The effects of fix length and flexible length parameters on the aerodynamic and flight dynamics of guided missile have been studied. A code has been developed to solve full Navier-Stokes equations using finite volume and modified Baldwin-Lomax turbulence model. Multi-block technique is also used to solve main body and fin parts flow field. Further, a 3 degree of freedom code has been developed to compare planar flight dynamic of missiles. It is found that missiles with bigger lengths for fix and flexible parts show more aerodynamic maneuverability, but drag force grows concurrently. Flight dynamic analysis shows that drag effect is negligible and aerodynamic maneuverability analysis is compatible with flight maneuverability.

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The flight dynamics of nine configurations of supersonic continuous deflectable nose guided missiles have been investigated. The studied configurations consist of a spherical nose tip, a tangent ogive, a set of stabilizing tail fins and a cylindrical body that its mid-section is flexible to form an arc of a circle. So the cylindrical body consists of a fix part in vicinity of nose, middle flexible part and main body with stabilizers. The effects of fix length and flexible length parameters on the flight dynamics of surface to surface, antiaircraft and antimissile missiles have been studied. A code has been developed to solve full Navier-Stokes equations using finite volume and modified Baldwin-Lomax turbulence model. Further, a 3 degree of freedom code has been developed to compare planar flight dynamics of missiles. This code consists of a guidance subroutine based on pure persuit law. The results show that even increase of fixed and flexible lengths enhance the maneuverability of the missile, but in some scenarios this can lead to increased flight time and more errors in the target engagement. Deflected nose relocates mass center away from the axis and a thrust vector torque is created. Study of surface to surface scenario shows that this torque improves accuracy of targeting and the ability of target dislocation. In air defense missiles, increase of Fix and Flex variables, will extend the limits of allowable firing angle. However, a heavy nose increases the role of thrust torque and subsequently decreases the role of nose geometry.

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Wind turbines are subject to various unsteady aerodynamic effects. This includes the wind gust and the change of wind direction. In this work, the aeroelastic behavior of a reference horizontal axis wind turbine has been investigated under different wind gusts and yaw conditions. Unsteady blade element momentum (UBEM) theory and Euler-Bernoulli beam assumption were used to rotor power estimations. To take into account the time delay in aerodynamic loads due to a sudden change in inflow conditions, a dynamic wake model was implemented. The ONERA dynamic stall model was coupled into the BEM theory to improve the aerodynamic loads prediction in the unsteady inflow and yaw conditions. To verify this method, the results in the case of steady-state are compared with the NREL 5 MW wind turbine and in the unsteady case are compared with the Tjaereborg test turbine. The results indicate that sudden change in wind speed causes sharp fluctuations in terms of elastic torsion of the blade and other parameters such as rotor power. Increasing in wind gradient can leads to increasing time delay to a new equilibrium. The increase in yaw angle can be contributed to the rotor power and the periodic loads reduction. The method presented here may facilitate improvements in the controller design for wind turbines.

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In this research experimental results of 60 degree delta wing airplane that conducted in National Low Speed Wind Tunnel is presented. The wind tunnel is closed type has an opened test section that its dimensions are 2.8 m × 2.2 m. Tests Reynolds number is beyond 1.5 million that achievement of this Reynolds number at low speed is unique in the country. Ground effect is measured using a fixed plane that its height is variable. Tests are conducted at the different height and aerodynamics forces and moments are measured using a sting type six component strain gaged balance. The tests results showed that the maximum lift coefficients increased from 1.29 to 1.38 due to presence of the ground plane. The lift coefficient due to ground plane in all range of angle of attack is increased and induced drag coefficient is decreased and consequently, the overall aerodynamics efficiency (lift to drag ratio) is increased from 8 to 14.5. When the distance between model and ground plane is less than half of the wing span, lift curve slope is increased in high rate from 2.66 per radian to 3.11 per radian. Decreasing this distance is caused the aerodynamic center is shifted backward and consequently longitudinal static stability is increased. Consequently presence of ground plane is caused increasing of airplane static stability.

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In this paper, a hypersonic inlet for operating at Mach 5.0 is designed and analyzed numerically. The main axis of this study is a series of three-dimensional simulations with the accuracy of 10E-06 which are applied to determine the effects of the highly developed boundary layer on the performance of inlet for three different study cases. The basic inlet concept is designed by integration of double ramp compression surface and inlet duct which can reduce the free-stream Mach number to the range of 2.0. The most important factor that it affects the performance of the hypersonic inlet system, is the developed entropy layer on the fuselage of the flight vehicle. Ingestion of this layer results in thermal gradients and pressure recovery losses. The bow shocks at the nose and the leading edges are the main sources of this low kinetic energy layer. Using the k-ω turbulence model in the numerical simulations have resulted in a reliable estimation of the boundary layer. In the current context, shock structures, shock-boundary layer interactions, flow quality at the end of the diffuser and also the effects of using sidewalls on the performance of the hypersonic inlet are the main goals of the simulations and the related results are summarized

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In recent years, Aerodynamic analysis of automobiles became one of the most important parameters which affect the power of the companies to be present in world markets. Therefore, they can be considered as one of the most important factors in aerodynamic design of vehicles. The formation of the vortex and consequently the pressure drop in the rear of the vehicle can increase the aerodynamic forces.
This paper investigates the methods for reduction of the vortices volume in the rear part of a sedan type vehicle by changing in geometry of the vehicles
. For this purpose, firstly in order to choosing the appropriate turbulence model and 3D simulation of incompressible flow around the Ahmed model (which its experimental results are available) was simulated using computational fluid dynamics. Then, the values of aerodynamic coefficients of a car model were studied by adding spoiler and creating curvature at its lateral surfaces. The results of this study indicated that the vortex volume formed at the rear of the vehicle can be simulated more precisely by using the Boundary-layer mesh around the model and analyzing the flow using the DES-SSTK-ω turbulence model Relative to the model K-ω-SST. Additionally, simultaneously use of the spoiler and the curvature of the lateral surfaces reduce fuel consumption and increase the stability of the vehicle due to a 26.3 % reduction in rear drag coefficient and a 5.2 % reduction in the lift coefficient, with respect to the simple car model.

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