نشریه دانش و فناوری هوافضا
سال دوازدهم شماره 1 (بهار و تابستان 1402)

In this study, a new approach is proposed to design low-thrust trajectory in the preliminary design phase for GEO to Halo transfer in the three-body problem. In this method, the approximate and near optimal trajectory is conducted through guessing thrust angles. In this approach, the in-plane and out of plane thrust angles are considered as finite Fourier series with unknown coefficients. The coefficients are calculated by an optimization algorithm with the aim of minimizing mission duration. Since the number of the trajectory revolutions is unknown, this parameter is also considered as a discrete decision variable of optimization algorithm. Due to the existence of continues and discrete decision variables, discrete particle swarm optimization algorithm is employed to solve the problem. The advantages of this method include: simplicity of execution and low volume of mathematical computation, considering satellite mass, do not assume special restriction for thrust angle such as tangential thrust and determination the near optimal mission duration. . In order to evaluate the proposed method, Trajectory design of GEO to Halo is performed for several level of thrust. Results indicate that this approach determines the number of trajectory revolutions, fuel consumed, near optimal duration and trajectory of the mission with high accuracy.

, the missile guidance and control system consists of three subsystems: navigation, guidance, and control. The task of these sub-systems is to calculate the deviation of the guided vehicle from the desired path so as to determine the appropriate movement or acceleration to compensate for the deviation. In the traditional methods , each of the guidance and control subsystems is designed separately, a. In the integrated guidance and control approach, the guidance law is developed separately and tested under the assumption of ideal autopilot. The autopilot is also designed independently and is tested under the assumption of an ideal guidance law. This paper describes the process of designing and simulating the performance of the optimal neural controller, which was created in order to guide the missile in a two-dimensional problem of minimizing the collision time and the distance from the target. In the design of the optimal neural controller, first the classical optimal neural controller (MLP) neural networks, the identifier, and the controller was designed and through simulation it was shown that the performance of this controller is not satisfactory. Therefore, by replacing the estimator MLP networks and controller with the deep type network, along with the use of the concepts of reinforcement learning, a quite improved performance was demonstrated through simulation. In this research, the integrated rocket model was made by integrating deep learning neural network with optimization algorithms, and the use of neural network control and optimization algorithms increased collision accuracy and reduced flight time.

In a holistic design; Creating a common language between different engineering, taking in to account the designer/customer's interests and preferences, removing potential shortcomings of common design methodologies and taking full advantage of the benefits of MDO approaches are the most important factors for achieving logical and realistic results. In this paper, a novel Comprehensive Preference-based Design approach is presented which attempts to achieve subjective attributes that are defined in the concept of maximization of designer/customer's satisfaction in addition to objective goals which are formulated in the form of minimization of a performance criterion in a two-phase structure using two nested optimizers. In the first phase of CPD, using the concept of satisfaction, the subjective preferences of the designer/customer are defined in terms of fuzzy relationships and operators in the form of demands and desires. Whereas the results of this phase are inaccurate, in the second phase, it is attempted to define a performance criterion and in order to achieve an optimal operational plan, attitude parameters and the compromises needed to meet the designer/customer's preferences are implemented. The methodology is utilized to design of a UAV with flight duration of 24 hours, 45 m/s of cruise speed at least, payload of 200 kg and less than 1 km take-off distance. To evaluate the results of CPD, the design process using MDO in AAO framework is also performed. Comparison of the results shows that despite the higher mass of UAV designed by CPD, overall satisfaction is higher and preferences have been satisfied.

In this paper, formation tracking control by maintaining a triangular shape for three quadrotors is discussed. Because of the nonlinear dynamic of the quadrotor and the presence of the external disturbance in the quadrotor dynamic model, a sliding mode controller was used for independent control of the quadrotor .To Formation Control of Quadrotors, a leader-follower decentralized formation control was applied. In such a way that the leader, follows a reference path and the followers use a formation controller to maintain a constant distance and an angle to the leader. First, the dynamic of the formation error is considered and then by using the integral slide mode controller, the formation error is reduced to zero. The simulation results obtained from the design of the sliding mode controller showed the occurrence of the Chatting phenomenon in the control commands, which was removed using the approximation of continuous functions. And the simulations obtained from this method showed the elimination of chattering in the control commands and the smooth tracking without oscillation of the desired path in the control commands. Also, the simulation results of formation control in the three missions with different paths, different formation geometry, and the number of different quadrotors indicate that the followers follow the leader and the success of formation control.

This article discusses the derivation of low-cost inertial navigation methods at stationary state in the navigation systems. To implement navigation algorithms, it needs to use an IMU (3 accelerometers and 3 gyroscopes). Thus, higher accuracy results higher cost. The methods used in this article do not need complete IMU sensors to calculate Euler angles, which includes an accelerometer and a rate gyro. Subsequently navigator hardware cost will decrease. For this purpose, mathematical equations are derived with multi-direction solution method, which the desired angle is extracted at first ,and then simulated as a hardware have done. specific values have been assumed for the Euler angles in the simulation process due to lack of access to field data, and the same angles are computed for two states (ideal and real sensors). By comparing the extended navigator algorithms, the best one (high accuracy and low cost) has been selected here in this paper.

Aerodynamic parameter estimation is an integral part of aerospace system design and life cycle process. Recent advances in computational power have allowed the use of online parameter estimation techniques in varied applications such as reconfigurable or adaptive control, system health monitoring, and fault tolerant control. The combined problem of state and parameter identification leads to a nonlinear filtering problem; furthermore, many aerospace systems are characterized by nonlinear models as well as noisy and biased sensor measurements. Extended Kalman filter (EKF) is a commonly used algorithm for recursive parameter identification due to its excellent filtering properties and is based on a first order approximation of the system dynamics. Recently, the unscented Kalman filter (UKF) has been proposed as a theoretically better alternative to the EKF in the field of nonlinear filtering and has received great attention in navigation, parameter estimation, and dual estimation problems. However, the use of UKF as a recursive parameter estimation tool for aerodynamic modeling is relatively unexplored. In this paper we compare the performance of three recursive parameter estimation algorithms for aerodynamic parameter estimation of two aircraft from real flight data.

Optimized structural architect of an Unmanned Lambda Wing Aerial Vehicle (ULWAV) during conceptual design has proven to be a significant challenge due to its unconventional configuration and no historical data for the structural architect. This paper aims to develop a new approach for structural design layout of new generation unconventional aircraft especially for ULWAV. The first step in this approach is to investigate the affecting parameters on structural architecture of these aircrafts then using of modal analysis to identification the behavior of these unconventional geometries for a general methodology development. The result of this modal analysis is apportion aircraft geometry to conventional substructures with already well-established defined architect. ULWAV structure architecture has been done by combining the structural architecture of sub-sections and with respect to requirements for systems installation, stealth design objective and packaging requirements. Here the CATIA software is used to optimize general architect with parametric model. To validate the results, the maximum static stress and maximum displacement of the structure were calculated using MSC / Patran software and compared with the values obtained from the optimization process. Dynamic analysis and buckling stability of structural composite shells are also controlled below. The architect methodology was applied to a case study of an ULWAV. The weight of the optimized structure has also been confirmed by typical case study researches.

This paper deals with a robust active vibration and non-singular fast terminal sliding mode control design for flexible spacecraft attitude maneuvers. First, the fully coupled nonlinear rigid-flexible dynamic model of the spacecraft in the three-axis maneuver is derived using Lagrange's equations in terms of quasi-coordinates. Then, the attitude control law is designed based on a fast non-singular terminal sliding surface, which leads to the zero convergence of attitude tracking and angular velocity errors in a finite time in the presence of external disturbances and parameter uncertainties. Next, the flexible panels' residual vibrations during and after the maneuver have been reduced exponentially using a robust active vibration control algorithm through piezoelectric sensor/actuator patches. It has been proven that this algorithm ensures the stability of the closed loop system and eliminates the need for conservative assumptions regarding uncertainties and external disturbances at the upper limit. The finite-time convergence of the closed-loop system with a hybrid control approach is proved by the Lyapunov stability theory. The numerical simulations using 4th order Runge-Kutta approach show the simultaneous utilization of the proposed attitude and vibration controllers' performance compared to the classical approaches for dynamical systems with structural flexibility.

Currently, the air fleet plays a significant role in the transportation of goods and passengers. Safety, quick movement and access to most important commercial and urban centers to the airport are the most important reasons for the importance of using the air fleet in the movement of goods and passengers. Wide-body passenger and heavy transport airplanes are used for a variety of missions, such as human and passenger movement, commercial goods and military equipments, etc. Their accidents are usually caused by one of human, environment, management, machine or a combination of two or more of them. In this article, using statistical analysis method and using SPSS statistical software, the main question titled: "What are the main factors affecting the occurrence of wide-body aircraft accidents?" has been studied. Also, suitable solutions to reduce the occurrence of air accidents of wide-body airplanes have also been suggested. In this research, the role and importance of the four factors of human, environment, management and machine in the occurrence of accidents and strategies to reduce the occurrence of accidents for these factors have been introduced. In this article, the main result has been obtained that the human role is the most important cause of wide-body aircraft accidents in the country.

In this paper, the finite element analysis of the free vibrations of functionally graded porous circular plate reinforced with graphene based on the first order shear deformation theory (FSDT) is presented for the first time. The governing equations are obtained using Hamilton's principle and the finite element method (FEM) is used to solve the governing equations of the sheet. The results of the present work were compared with previous studies and a good agreement between the results was observed. The effect of different parameters such as porosity distribution, porosity coefficient, different GPL patterns and weight percentage of graphene nanoparticles, types of boundary conditions and also the ratio of thickness to radius on the vibrations of the circular sheet was investigated. Next, while validating the analysis method, the results obtained from the numerical solution were compared and analyzed. In the results, it was found that the effect of weight percentage of graphene and different types of graphene patterns as well as support conditions in sheet vibrations is more than other cases.

The present study aims to investigate the vibrational behavior of composite conical shells reinforced by an anisogrid lattice structure, using the analytical method. For this purpose, the smeared method was employed to determine the equivalent stiffness contribution of the stiffeners. In this approach, the smeared method was used to superimpose the stiffness contribution of the stiffeners with those of the shell to obtain the equivalent stiffness of the whole structure. The stiffeners were modeled as beams that can support shear forces and bending moments along with the axial forces. The governing partial differential equations of the problem are derived by applying Hamilton's principle and based on the first-order shear deformation theory, and then they are converted into a group of algebraic equations by using the differential quadrature method. Then, eigenvalue problem, and as a result, natural frequencies are calculated. In order to validate the results, comparisons of the present results with those of other studies are performed. Then. the effect of different parameters such as semi-vertex angle, circumferential wave number and the effect of different boundary conditions on the natural frequencies of the system has been evaluated.

In this research, authors attempt to investigate the significant and influential characteristics behaviors of the conical nose in supersonic flow as well as discussing the results of using Multi-Row disks and their effects to recognize the behavior of this nose. A cone nose produces low drag and high aerodynamic heat due to sharp tip in flight conditions, which makes it problematic at high speed Mach numbers. Thus, new modifications with the utility of adding Multi-Row disks in the design of this cone nose allows the nose to efficiently reduce the produced aerodynamic heat as well as preserving its nature to produce low drag. Results show that the using 12 disks led approximately 30% reduction in drag coefficient rather than reference nose without disks. It also significantly reduces the static temperature produced on the nose surface. Therefore, this type of nose improves its aerodynamic characteristics that can advise to be used at high speed and raises the possibility of reusing this kind of nose instead of the strict conical one.

In this study, the effect of blowing angle on coefficients of aerodynamic of NACA0012 airfoil at Re=4×106 has been investigated. In the present research, the effect of three-dimensional blowing jet on the aerodynamic performance of the wing has been considered. The changes of lift and drag coefficients were studied by using the blowing near the leading edge of the airfoil for angles of attack of 12, 14, 16, 18 and 20 degrees. Due to the application of blowing through circular holes, a three-dimensional solution was used for the purpose of analysis and study, which will be computationally expensive. The analysis carried out in this case was performed with the assumption of incompressible and steady flow around a three-dimensional wing section in Fluent software. The results showed that the blowing has a negligible effect on the lift and drag coefficients at angles of attack less than 14 degrees. It is for this reason that only the effects of the blowing at high angles of attack are considered. The greatest increase in the lift coefficient and the greatest decrease in the drag coefficient occurs at the angle of attack of 16 degrees, which is the stall angle. The results showed that as the jet angle increases, the aerodynamic performance decreases.

In this study, the performance of the air distribution system in the cabin of a single-aisle passenger plane with 42 passengers seated in two rows of three has been analyzed. Computational fluid dynamics method has been used to investigate the air distribution in the cabin. In this numerical analysis, the effect of changing the relative humidity inside the cabin, changing the speed of air entering the cabin and changing the way of air distribution inside the cabin has been investigated. In order to determine the thermal comfort in the cabin, two parameters including the average number of predicted votes and the percentage of predicted dissatisfaction have been investigated. The results show that with the increase of relative humidity from 0 to 10%, changes in static pressure do not cause major changes in passenger comfort. Also, at a speed of 1 m/s for air entering the cabin, due to better air circulation in the head and other organs, the passengers feel more thermal comfort and less static pressure changes on their head and body. Finally, by changing the air distribution system from the mixing system to the displacement system, it was observed that the mixing air distribution system is better in terms of thermal comfort for passengers. However, in terms of static pressure in the head and body area of passengers, the displacement air distribution system is better.