نشریه دانش و فناوری هوافضا
سال دوازدهم شماره 2 (پاییز و زمستان 1402)

Numerical analysis of damage of large structures in real dimensions is a complex and time-consuming analysis. Therefore, it is important to use a simple and accurate damage model. In this research, the test and numerical analysis of damage in the composite wing structure of an airplane in real scale in bending loading was done. For a more complete review of the test results, extracting more data about the way of causing damage and also estimating the strength of the structure, the wing structure including spars and ribs, as well as the effect of fabric patches in the finite element software ABAQUS was modeled and subjected to damage analysis using USDFLD subroutine. After the initiation of damage, this subroutine reduces the mechanical properties suddenly. Examining the results of numerical and experimental analysis showed that the used damage model can accurately simulate the mechanical behavior of the wing structure. From the results, it was found that the numerical model and the composite wing structure have the same stiffness; but the strength of the analyzed wing is lower than the final strength observed in the bending test. Comparing the results of the strain gauges with the results of the analysis also shows that the modeling of the structure with the  layup details is done correctly and the damage model is implemented with appropriate accuracy in the software. Therefore, it seems that, this damage model can be used for damage analysis of real and large composite structures with appropriate speed and accuracy

In this research, using analytical method, sound transmission loss in fiber metal laminate (FML) composite cylinder is studied. For this purpose, an infinitely long cylindrical shell composed of FML is subjected to an oblique plane wave. considering the effects of shear forces in the sound transmission loss, specifically at the higher frequencies, the FSDT theory is used to solve the governing equation of the motion. The equation of motion of the shell are obtained by D'Alembert principle. The equations are solved simultaneously using acoustic pressures, boundary conditions and infinite series that converged with a special algorithm, shell equations and acoustic relations. The values of sound transmission loss from numerical solution are validated with other researchers. This study shows that using FML cylinder instead of composite cylinder, the coincidence frequency increases 18 percent. Also in mass control region transmission loss increases 9 percent. It is demonstrated that with constant volume fraction of metal, locating aluminum lamina in farther distance from neutral axis, ring and coincidence frequency changes 11.4 and 18.7 respectively. In addition, results indicate that by increasing of 20 percent of metal volume fraction (MVF), in a specified frequency, transmission loss can be increased 28 percent.

The strong voltage difference between two electrodes with different thicknesses causes the corona discharge phenomenon. Today, corona discharge has various applications, including use in thrusters. In the last decade, the use of corona plasma actuators to control flow and generate thrust has received much attention. In the present study. COMSOL software has been used to perform the simulation. First, the simulation results are validated in the wire- cylinder configuration and then the corona discharge in the wire-airfoil configuration is simulated. In the present study, the effect of the distance between the two electrodes on the airflow characteristics due to corona discharge such as electric current, thrust, fluid flow velocity and electrical power consumption is studied numerically. The results show that if the distance between the two electrodes increases, the electric current and the electric wind speed decrease, but the efficiency and power consumption increase. The rate of change of thrust increases by 15% for each kilo volt increase in applied voltage.

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.

In this article, the flow separation control by the vortex generator jet actuator on an airfoil with NASA SC(2)-0714 cross-section in the flow with Reynolds 1×106 has been numerically investigated. The pulsed jet has a frequency of 40 Hz and a duty cycle of 50%. The simulation has been done at the angles of attack of 14, 16 and 18 degrees where the flow separation on the airfoil surface starts and then comes to a stall condition. In this research, in order to better understand how the pulsed vortex generator jet works, the effects of the unsteady jet on the instantaneous and time-averaged flow characteristics are extracted and the vortices pattern of the flow on the airfoil surface has been investigated. The results indicate that the actuation effectively delays the flow separation at the angles of attack of 14 and 16 degrees. As a result, the aerodynamic efficiency of the wing section increases by 14.8% and 33.2%, respectively. At the angle of attack of 18 degrees, which is in the post-stall condition, the actuation eliminates the massive flow separation and the flow attaches to the surface in a significant part of the airfoil chord.

In this study, using Fluent version 2021 software, to investigate the effect of volume fraction, particle diameter in the Reynolds number range of 10,000 to 100,000 using finite volume method and Eulerian-Lagrangian approach for two-dimensional and three-dimensional geometry. The next one is done in a 1 meter tube. Also, the application of thermophoresis and Brownian force and its effect on Nusselt number results were investigated. The data related to all three modes of thermophores and Brownian force and in the absence of these two forces up to Reynolds 60 thousand were in good approximation with the experimental results and the effect of the forces and the difference of the respective Nusselt numbers from Reynolds above 60 thousand can be seen. The results show that with the addition of nanoparticles, the thermal performance coefficient is increased, and the size of the diameter of the nanoparticles and the volume fraction have a significant effect on this performance. The results were validated with experimental values and showed good agreement.

The wireless communication of unmanned aerial vehicles (UAVs) has attracted a lot of interest in military and civilian applications for reasons such as the possibility of high mobility, low cost, deployment based on needs, and the efficiency of direct air-to-ground channels. However, these advantages make UAV wireless systems vulnerable to both passive and active attacks. In recent years, with the increase in the computing power of computers and as a result of the increase in security threats, a new and efficient security solution called physical layer security has received much attention. In this field, one of the most important methods is the key generation from wireless channel characteristics. In this paper, by focusing on the key generation from the channel phase, two methods of generating the key from the channel phase are introduced: 1) the first method, using the phase difference of a sinusoidal signal with two frequencies and 2) the second method, choosing the random initial phase for a sinusoidal signal with one frequency. The simulation results show that key generation based on the second method performs better than the first method from the perspective of active attacks (man-in-the-middle attack and jammer attack). Furthermore, we study the second method from the point of view of key disclosure resion. We will see that by implementing the idea of channel probing on multiple frequencies, the percentage of key disclosure region will decrease. The results obtained provide useful engineering insights for the design and optimization of a secure link for UAV communications.

In current study, an integrated guidance and control system is assessed for a projectile having dual spin motion in the cascade control structure. In this way, a nonlinear seven degrees of freedom model is used and coupling effects of pitch and yaw channels is taken into consideration. Based on cascaded control structure, a three-loops controller is designed, which in inner and middle loops the dynamic inversion method is used, and in the outer loop, an adaptive model based on PID controller is designed. The dynamic pressure contribution in acceleration command is also taken into account. The performance of the designed guidance and control system is assessed in presence of uncertainties via Monte Carlo statistical simulations. The results of suggested algorithm are compared with that of a cascade control having a classic PID controller in its outer loop. The results show the superior performance of the proposed algorithm in achieving the target and control effort.

The presence of the actuator saturation in a three-loop autopilot of a flying vehicle arising from the windup phenomena leads to the performance degradation. In this paper, a dynamic anti-windup and a static anti-windup compensators are proposed. The dynamic compensator, available in some references for systems with output feedback is designed to make the controller states of the saturated system arbitrarily close to the controller states of the unsaturated system. In this paper, this compensator is developed for controllers similar to three-loop autopilot of a flying vehicle, in which, the controller is a feedback of both plant output and plant states. In order to compare with the dynamic compensator, a static compensator is also designed with a new approach in such a way that the controller output tracks the output of saturation function. With this requirement, it is expected that no windup will occur. These two compensators have been applied on the autopilot of an aerodynamic tail-controlled flying vehicle in two stable and unstable conditions and their performances have been compared to each other in terms of control performance. The results show the better performance of the static compensator compared to the dynamic compensator in tracking the desired command in both operating conditions. Moreover, through simulation of flying vehicle guidance system, it is shown that using anti-windup leads to the less miss distances. Especially, using the static compensator leads to reduction of the miss distance by about 50% in unstable conditions compared to the use of the dynamic compensator.

In this article, an analytical method for detecting the need for pressure feedback is introduced in controlling the position of the hydraulic actuator’s piston. The geometric location of the roots is used to do this. In this method, with the appropriate conversion function for the actuator, the effect of changing the position and pressure feedback parameters on the location of the close loop system poles and thus its bandwidth and relative damping is examined and based on it. First, the use of pressure feedback or no need for it is specified, and in the second, a guideline for designing pressure feedback is provided when necessary. By presenting this analytical method, it is first saved at the costs of preparing pressure sensors, locating and installing them, as well as pressure feedback parameters in unnecessary times, and secondly, a specified instruction for the design of the pressure feedback on times of use is extracted from it. Finally, by introducing and simulating three samples with dynamic characteristics and different conditions, the accuracy of the diagnosis and design method presented to control these types of operators is described.

Generally, the guidance and the control are designed as separate modules, which are suitable for many common and normal applications, but in order to achieve less error in deploying the satellite into the orbit increase the accuracy in performing the mission, we need to develop integrated Guidance and Control algorithms (IGC). By using such algorithms, it is possible to design the closed loop guidance for the last stage of the launch vehicle in vacuum condition. Examples of this are iterative guidance mode (IGM) and power explicit guidance (PEG). In this paper, the guidance of the final phase along with the guidance of the first and second stages of a typical solid rocket launch vehicle is modeled and simulated. The results of this article show that the IGM guidance can be well compatible with the Pitch Program guidance of the previous steps and successfully complete the deployment in the orbit.

With the increasing development of space technology applications and the consequent increase in the requirements of space missions, many satellites have changed their nature from rigid to flexible structures. Attitude control of flexible satellites is more challenging than rigid satellites. This survey examines the challenges and architectures of flexible satellites and analyzes their dynamic behavior. Moreover, a comprehensive review of various attitude control approaches which have been proposed by researchers in recent years for flexible satellites has been done and the advantages and disadvantages of each of them are compared and evaluated to other approaches. Finally, based on the conducted survey, some recommendations have been made for designing an appropriate attitude controller strategy for flexible satellites.

The purpose of this article is to optimize the power supply subsystem of the satellite based on mass and power production. For this purpose, the satellite orbit is to determine the time of placement in the shadow in Satellite Tool kit (STK) software, and the design of the satellite energy supply subsystem of existing analytical equations are done, and these main design parameters have been optimized to achieve the minimum weight of system. Then, the beginning of the target satellite, the Orsted satellite, the orbital and mission operational modes of this satellite has been compiled and the statistical information of its power supply subsystem has also been determined. Then, the design of the power supply sub-system is based on the parametric analysis and processing of statistical information and also based on the simulation and orbital analysis of the satellite. Finally, in order to model the design problem, the objective functions of design and optimization, the weight of the power supply subsystem has been considered, and the main design parameters affecting the objective function are capacity, depth of discharge, efficiency and density of the battery and the area of ​​the arrays. It is worth mentioning that the type of battery itself is also very effective on the objective function, but considering the type of satellite battery considered (nickel-cadmium), the variable battery type is not considered. Finally, this problem has been optimized using genetic algorithm. The obtained results show the accuracy of the method and implementation compared to the values ​​of the existing satellite power supply subsystem.

Solution of the static attitude determination of the satellite leads to the Wahba problem. The Wahba problem uses a set of at least two independent sensor measurements and reference vectors. These input vectors are not accurate due to sensor noises, misalignment, and low-order approximations. But these uncertainties do not view in the classic Wahba problem directly. Hence, the estimation error of the Wahba problem depends on the accuracy of the input vectors. In this paper, modeling error, measurement noise, and biases are proposed unknown but bounded. These errors are modeled using interval analysis. The innovations of this research are considering the uncertainties in the input vectors in the Wahba problem using interval arithmetic and transforming the solution of the attitude determination problem from a single-objective problem to a multi-objective robust optimization problem. Then the multi-objective problem is optimized using an NSGA solver. The results indicate a lower attitude estimation error of the proposed method in attitude determination of the satellite.