mohammad hassan djavareshkian

In this research, one of the widely used control systems of tailless aircraft, named Crow Flap, has been simulated by numerical method and the appropriate opening direction of this system has been investigated. By creating a differential drag on both sides of the aircraft, this system creates the necessary yawing moment and directional control. This system consists of two moving surfaces that are deflected against each other to neutralize each other's force and produce drag. One of the important issues in control systems is the reduction of disturbing moments. Determining the suitable opening direction with the least drag and disturbing moment and increasing the useful moments is the aim of this research. For this purpose, the measurement of these moments and forces are compared in two states of normal and reverse-opening. The UAV under investigation is a lambda-shaped UAV named Swing. Experiments have been carried out at AOAs of 0 to 12° for three opening angles. The created Mesh is unstructured. In this simulation, the continuity, momentum, and scalar equations have been solved by the Simple C algorithm after discretization with the finite volume method, and the turbulence model (K-ω-SST) has also been used. In this research, the opposite opening direction of the Crow Flap, unlike the usual opening direction mentioned in some articles, had a better result, so that at high AOAs, this opening direction has between 5 and 32% less drag and it has produced 10.5 to 35% more Yawing moment along with reducing disturbing moments.

An unmanned aircraft (UAV) is a type of aircraft that operates without the need for a human passenger. These planes are able to fly automatically, without the need for remote control, and are mostly controlled by computers and sensors. UAVs are used in many industries and have various applications, including communications, surveillance, aerial photography and videography, security and border surveillance, search and rescue, scientific and environmental studies, agriculture, as well as the military industry. The advantages of using UAVs include high efficiency, access to difficult and dangerous areas, and cost reduction. However, the use of UAVs also comes with challenges and problems. One of the most important problems is the risk of losing control, aerodynamic problems, and running out of fuel at different angles of attack. Improving aerodynamic coefficients can help increase flight safety and reduce risks related to losing control of the aircraft. In this research, we examine the elements, changing the shapes created on the airplane wing, and applying control surfaces to improve the aerodynamic coefficients.

This research delves into the intricate realm of supersonic inlet design for ramjet engines, honing in on the critical aerodynamic considerations and optimization of performance factors. At Mach 2.5, the study meticulously scrutinizes pivotal design parameters, including the placement and number of inclined shocks, cowl-lip positioning, throat area, spike location, and diffuser length. Computational fluid dynamics simulations are harnessed to unravel the intricate flow dynamics and assess the proposed inlet geometry's performance.The findings reveal a nuanced relationship between back pressure and shock wave positioning, where increasing back pressure initiates a shift in the shock wave, impacting the flow state. The paper delineates this transition, emphasizing the pivotal back pressure range of 300,000 to 350,000 pascals, where optimal shock wave alignment corresponds with design parameters, achieving a supercritical state.However, elevating back pressure beyond this range triggers a sub-critical state and mass flow overflow as the shock exits the throat.the study explores various performance metrics, encompassing drag coefficient, distortion coefficient, mass flow ratio and total pressure recovery under varying back pressure conditions. The outcomes underscore the merits of higher back pressures, which mitigate drag coefficient and distortion while amplifying TPR.In the sub-critical state, MFR diminishes due to shock wave displacement beyond the intake opening.This research illuminates the intricate dance of aerodynamics within ramjet engine inlets and underscores the paramount significance of optimizing inlet geometry to unlock heightened performance. It effectively encapsulates the essence of the full article, enticing readers to embark on a deeper exploration of this crucial area of aerospace engineering.

The research investigates the growth of ice on the wing of a UAV, both with and without a winglet, and its effects on aerodynamic coefficients at low Reynolds numbers using a numerical method. The wing configuration is rectangular, employing the NACA0020 airfoil cross-section. Simulations were performed using a numerical method based on the finite volume method, a pressure-based algorithm, and second-order upwind scheme for convective flux calculation. Turbulent flow is modeled using the Spalart-Allmaras turbulence model. For ice modeling, the FENSAP-ICE commercial software, a modular ice simulation system, was utilized. The iterative ice accumulation simulation process includes successive calculations of airflow, water droplet paths, collection efficiency, and heat transfer balance to determine the shape of the accumulated ice. Calculations were conducted at Reynolds numbers of at an angle of attack of 10 degrees. Results indicate that the ice profile formed on the leading edge of the airfoil generates a swirling flow in this region due to depressions above and below the ice mass. This swirling flow increases the back pressure coefficient and decreases the drag coefficient compared to the wing without ice. Adding a winglet alters the airflow and reduces air resistance by controlling the induced vortices from the wingtip, thereby increasing lift and reducing drag. Consequently, the airflow speed increases, reducing ice accumulation and enhancing the aerodynamic efficiency of the wing.

This research investigates the effect of fence on wingtip vortices and control surfaces in a bird-like aircraft using numerical methods. In designing and placing the fence, average dimensions extracted from wing root vortices at different angles of attack have been used. In addition, the fence with specified dimensions have been installed at three heights and three different positions along the wing (the length of the winglet is equal to the average length of the vortex in that part, and the height of the winglet is 30% of the diameter of the vortex in that part) and have been examined at angles of attack ranging from 7 to 16 degrees. The next stage of the study is the optimal design of the dimensions of the control surfaces using a single objective optimization method. The aim of this research is to achieve the best possible design that converges to an optimal solution with minimum time and cost. The design of control surfaces is carried out at three points based on the dimensions of the wing root vortex using numerical methods. However, Computational Fluid Dynamics (CFD) analysis requires a lot of computational time.

In this research, wing fences have been designed using the parametric study method. Here, by deflecting the vortices of the wing apex, the excess rolling coefficient produced by the split drag rudder system has been reduced. This control system is used to generate yawing moment in flying wing UAVs. The size of the fences is selected based on the dimensions of the vortex at different AOA. In this research, the survival equations are discretized by the finite volume method, then the discretized algebraic equations are solved by the Simple C algorithm. In this simulation, the K-W-SST two-equation model is used to model the turbulent flow. The UAV used in this experiment is a lambda-shaped flying aircraft. The produced fence was installed at three different heights and three different positions along the wing, which were investigated at high AOA. The results show that the use of fences in all AOA will not be suitable from an aerodynamic point of view; But finally, the most optimal location of the fence will be introduced to reduce the disturbing rolling moment coefficient.

Numerous studies have demonstrated that human factors are the primary cause of deadly air incidents and accidents involving passenger flights. Human error is defined as mistakes and flaws in the way a human component of a system performs a predetermined action or refrains from performing an activity that is forbidden or that needs to be completed within a given amount of time, with a specific degree of precision, or both. Vision error, measurement units, radio communication, language, aircraft warning system design, psychological stress, pilot selection and qualification, pilot flight qualification, fatigue, aging, human performance, flight crew coordination, modifications to the Course of Cadin course, and human factors in aircraft design are the typical categories into which such errors are divided. Based on international norms and indications, the investigation's findings demonstrate that the Civil Aviation Organization's supervision in the sensitive area of ISI has been appropriately followed. This matter is significant because, since 2010, the Civil Aviation Organization has complied with its standards optimally based on international audits. This is because, among domestic experts, compliance with international standards has always been a top concern in increasing safety and productivity in this field. However, determining if this issue is sufficient to establish safety and prevent accidents is crucial. In this research, we examine the risk method for flight safety, the performance evaluation index of safety, security, and flight safety by passengers, and the most important golden tips to save from plane crashes.

This research explores and compares the AUSM scheme family based on compressible, steady, viscous, and inviscid axisymmetric flows in a finite-volume method-based and unstructured data storage grids code. When the effects of side-velocity are taken into account, axisymmetric flows can be considered a three-dimensional problem in the longitudinal plane; as a result, there is a considerably decreased number of computations required compared to computations in three dimensions. The most important and latest modifications of the AUSM-family were developed to identify more efficient methods in the AUSM- family in terms of accurate prediction of the axisymmetric flow field in internal and external axisymmetric flows, viscous (inviscid), and high-speed flows with shock waves characteristics. The novelty of this investigation is the assessment and comparison done on the AUSM-family in resolving the compressible axisymmetric flow field, which has received less attention in prior studies. The studies determined that the AUSM+M method performs better against a strong shock wave than other methods investigated in this research. According to the modifications made in this scheme, unlike other methods, there are no wiggles in the regions mentioned earlier. Furthermore, it is discovered that the AUSM+M method had a higher rate of convergence than the AUSM+ and SLAU approaches. In addition, the AUSM+M scheme is distinctive from other techniques in viscous flows because it produces the minimum kinetic energy dissipation rate and fewer shock anomalies in the shock wave region.

In this research, the placement of the split drag control system along the length of the UAV wing and its effect on the aerodynamic coefficients are numerically investigated. This control system consists of two plates on top of each other, which, when opened, creates a pressure drag in one wing. This system is used to create a yawing moment in flying wing airplanes. Flying wing airplanes have a high sensitivity for determining the location of control surfaces due to the presence of the swept back angle in the wings and the formation of the wing apex vortex at high angles of attack in this type of configuration. Here, for the installation and positioning of the split drag control system, from a static point of view, it is necessary to install the moving surfaces of the split drag at the end of the wing, because the maximum moment arm will be in this part, which causes the production of the maximum yawing moment; However, from the aerodynamic point of view, the placement of the control surface in this range always has disadvantages due to the existence of the wing tip vortex and the wing apex vortices. Therefore, here it has been trying to place the split drag system in 3 different opening angles in 3 longitudinal positions relative to the tip of the wing and check the resulting moments in different angles of attack. The aim of this research is to increase the yawing moment and decrease the rolling moment coefficients.

In this research, the split drag system at different AOA for a flying wing aircraft has been simulated and optimized by a numerical method. The split drag system provides vertical axis control by creating asymmetric drag between the right and left wings. The aircraft under study is a lambda shape aircraft with a swept-back angle of 56. The split drag control system is made up of two surfaces on top of each other, by opening in the opposite direction on one side of the aircraft, it creates the drag necessary to produce yawing moment. Their installation position is at the tip of the wings and on the trailing edge. When using split drag, in addition to the yawing moment, a disturbing rolling moment is created, which is caused by the drag difference between the upper and lower surface of this system, and the reason for this is the change in the AOA of the aircraft. Asymmetric opening of the surfaces can reduce the induced roll to zero and in some cases to a minimum. The test was carried out in AOA of 0 to 12ᵒ for drag split opening angles of 10, 20, and 30ᵒ. Calculations based on equations (RANS) are discretized with the finite volume method. The obtained results show how much to add to the angle of the split drag surfaces depending on the AOA in order to find the most optimal mode to neutralize the roll, and finally, the optimized diagrams of this system are obtained.

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.

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.

Optimization of the height and length distribution of roughness elements as an effective passive flow control tool was investigated in the current study. The purpose of this paper was to investigate the roughness element's effect and its location on upstream of the laminar separation bubble from phase portrait point of view. Consequently, the effect of the roughness element features on the bubble's behavior is considered on the vortices behind the NASA-LS0417 cross-section at the pre-stall angles. The consequences express that the distribution of roughness in the appropriate dimensions and location could contribute to increasing the performance of the aerofoil and the interaction of vortices produced by roughness elements with shear layers on the suction side. It is worth noting that due to the small size and low velocity of the flight of MAVs, the formation of the well-known phenomenon of laminar separation bubble is almost imminent. Since this phenomenon greatly affects the aerodynamic performance and patterns of the Phase portrait circuit, its recognition, investigation and control can be a key parameter. In the meantime, the results show the emergence of nested loops from the Phase portrait circuit due to flow arrangement changes. Also, the distribution of roughness in the appropriate dimensions and location can be recognized as a factor that helps to increase the performance of the aerofoil, and can be the basis of the work of MAV designers.

In the present study, the ground effect on the dynamics stall of the flapping wing with the bending deflection angle is experimentally investigated in the forward flight.For this purpose, first, a bending deflection mechanism and the installed facilities for the wind tunnel have been designed and fabricated.Then the simple and bending flapping wings are experimentally examined for h/c=1,1.5, flapping frequency 3.5 Hz and different angles of attack from 00 to 22.50 with velocity 3 m/s. Also, to investigate the effect of flapping frequency and distance from surface on lift force, thrust force and loading power of the simple and bending flapping wings with no angle of attack tested for h/c=1,1.5,2, different flapping frequency from0 Hz to 5 Hz and velocity 3 m/s. Results indicated that decreasing the distance from the surface, dynamic stall of the simple flapping wing occurs at lower angles of attack compared to the bending flapping wing. More specifically, in the minimum distance from the surface (h/c = 1), the stall angle of the simple flapping wing and bending flapping wing (with bending deflection angle of 1070) takes place at 12.50 and 150, respectively. The performance of the bending flapping wing is generally better than the simple flapping wing. Besides, by enhancing flapping frequency and decreasing the distance from the surface, aerodynamic forces and loading power increase.

In this research three different models of rotational force are evaluated in the study of a hovering fruit fly inspired wing, with combined flapping and pitching motions, using a quasi-static simulation by the blade element method, and the results achieved thereby, are compared with previously published results. Then these models are compared with each other, and the model which has the lowest relative error is introduced. The traditional rotational force model, which depends on the wing's translational velocity, does not consider any rotational force at the beginning and end of the stroke. The new rotational force models which are designed to accurately predict this force also consider a second component for this force which is due solely to wing pitching. In this research, the instantaneous and mean coefficients in the blade element method are compared with published computational fluid dynamics and experimental results. The models are compared in terms of the error in predicting the maximum point on the instantaneous force coefficients curve. The root-mean-square error analysis shows that one of the rotational force models, which includes the two force components described by the Kutta-Joukowski theorem and the force due solely to wing pitching, has higher relative accuracy than other models and can be proposed for quasi-static simulations.

In this research, a different view of the study of the pattern and behavior of turbulent flow on the two full-span wings of baseline and sinusoidal leading-edge with periodic boundary condition by a numerical method is introduced. In this simulation, the Navier-Stokes equations are discretized by the finite volume method and solved using the improved turbulence model (IDDES). In other words, in the current study, to enhance the maneuverability of a fix-winged MAV, a passive flow control method was employed, which was inspired by the buoyancy fin of a special species of whale called a humpback. In order to study the performance of this type of infinite wing, the Reynolds number was considered equal to 140000. The results show extensive variations between these two types of full-span wings. The sinusoidal leading-edge full-span wing, in contrast to the baseline infinite wing, has severe fluctuations in pressure distribution and flow pattern at the pre-stall region. These variations are caused by the dominance of lateral flows over longitudinal flows on this type of infinite wing. Therefore, researching the patterns and analysis of vortices formed on this type of aerofoil as well as the associated frequencies is always a useful tool for understanding physics of flow on them, which will provide designers with a new perspective on flying objects.

The twist is one of the most important parameters in the design of the flying wing and tailless aircraft that causes eliminate some aerodynamic challenge at these categories of aircrafts. The present study was performed for an aerodynamic investigation of the geometrical twist at a subsonic flying wing and evaluate this parameter at different flight phases. The study geometry is a lambda-shaped flying wing that has a wing with a 56-degree sweepback. The twist angle applied on wingtips is washout, which is linearly distributed along the wingspan. The study is conducted in the framework of numerical simulation and based on solving Reynolds-Averaged Navier-Stokes (RANS) equations by finite volume method. The simulation process was performed after validation with experimental data, for twist angles of 0 and 6 degrees and range of attack angles of 5 to 20 degrees; also, to investigate the twist performance in the range of landing and take-off phase and cruise phase, studies have been performed in two different Reynolds numbers. The results show that by applying twist, the aerodynamic efficiency is improved at high angles of attack, but this characteristic will drop significantly at the zero-degree angle of attack. Also, by applying the twist, the conditions required for longitudinal stability are satisfied, and the pitch up phenomenon will be delayed.As speed increases, aerodynamic efficiency improves over a wide range of attack angles; also, aerodynamic efficiency changes due to twist increased, and twist will be more effective. Pitch moment analysis shows that as speed increment, the degree of stability will increase, and the pitch-up behavior will improve.

In this study, the effectiveness of three types of winglets such as Blended , Multi-tip, and Raked on a specific wing,is investigated by using a numerical method based on finite volume and pressure-based algorithms. In thisnumerical method, the Spalart-Allmaras turbulence model is used. In this simulation Reynolds number is 1.5×105and SD7032 airfoil is used for wing section by 4.6 aspect ratio. The stresses on the wing surface are calculated bythe wall functions, and the convective fluxes are computed by the second-order upwind accuracy. In this research,the specific airfoil is used for wing and winglet and the effectiveness on the aerodynamic performance of arectangular wing has been investigated. Evaluation of aerodynamic coefficients and flow physics have shown thatusing the same airfoil for wing and winglet with different angles of attack, It has increased the performance of theMulti-tip winglet by 3.1 percent compared to the case of using two airfoil and also for Blended and Multi-tipwinglets compared to the wing without winglet will increase an average aerodynamic performance by 10.5 and 10percent. But for Raked winglet, it has only a small effect on reducing the power of the vortex core.

The prediction of aerodynamic characteristics and smart flap of airfoil under the ground effect are carried out by the integration of computational fluid dynamics. Considering different types of beams، a parametric bending profile of a smart flap is designed. Cantilever beam with uniformly varying load with roller support at the free end is considered in here. A pressure-based implicit procedure is utilized to solve Navier-Stokes equations and a nonorthogonal mesh with collocated finite volume formulation is used for simulation. The boundedness criteria for this procedure is determined by Normalized Variable diagram (NVD) scheme. The procedure incorporates the eddy-viscosity turbulence model. The SIMPLE algorithm is applied for turbulent aerodynamic flows around the airfoil with smart and conventional flaps. The results of two flaps are compared for different flap length، flap angle and ground clearance. It is found that: 1- The pressure coefficient distribution in smart flap is smoother than conventional flap. 2- Lift-drag ratio in smart flap is higher than conventional flap. 3- The maximum lift-drag ratio is at flap angle 7. 5º. 4- The minimum ground clearance has the highest lift-drag ratio. 5- The increasing of flap length leads to the increasing lift-drag ratio.