نشریه دانش و فناوری هوافضا
سال دهم شماره 2 (پاییز و زمستان 1400)

Double-walled structures are widely used in various industries such as aerospace, marine, and automotive. Therefore, in this paper, the sound transmission in these structures is investigated. Due to the influence of rotation and shear parameters by increasing the thickness of the cylindrical shell, the Newton-based three-dimensional theory of elasticity is used. In order to solve the governing equation of motion for a cylinder, it is assumed that the displacement field is the sum of the gradients of a scalar potential field and the curl of a vector potential field. As a result, the shell motion equation becomes two separate wave equations, which solve the displacement field. To confirm the obtained equations, the present results are compared with the results of other researchers in this field that have been obtained with other theories such as classical shell theory. Finally, the effect of various parameters such as properties and thickness of the fluid layer, Mach number, and material of the cylinders are investigated. The results show that in double-walled structures with the air gap, the acoustic impedance (the speed of sound multiplied by density) of the fluid is the main and effective factor in sound control. Any fluid with more Acoustic impedance behaves better in sound control and improves sound transmission loss.

In this paper, the vibrations on the sample of a single-rail sled tester system with dampers are investigated. First, by extracting the vibration equations governing the problem coded in MATLAB software and by drawing the FFT diagram, the natural frequencies of the system are obtained. The first and second frequencies of the system are related to torsional displacement and transverse sled displacement, respectively. Due to the structural damping of the sled system, to extract the equivalent stiffness and equivalent damping values, the experimental test is used by the harmonic dynamic test device. Then, by simulating the sled model in ABAQUS software, modal analysis is performed and natural frequencies and mode shapes are extracted. Finally, by constructing a designed sled model and performing experimental modal analysis, the natural frequencies of the system are compared with the previous two methods and confirmed with an error of less than 9%.KeywordsSled tester system, natural frequency, modal analysis, equivalent stiffness, equivalent damping.

This paper deals with active vibration control (AVC) and structural health monitoring of a maneuvering flexible spacecraft with a cracked panel using piezoelectric (PZT) sensor/actuator patches and strain rate feedback (SRF) method. The cracked panel is modeled using the Euler-Bernoulli beam theory and the finite element method (FEM). The nonlinear equations of motion of the fully coupled rigid-flexible system are extracted using the Lagrangian formulation and solved with Newmark-β algorithm. Two approaches of structural health monitoring; first, trial and error, and second, maximum strain rates (online measurement) along with AVC (applying control signals based on the maximum values of strain rates to the predefined number, but unknown locations of PZT actuators), are performed. The strain rates change directly with mission conditions and crack locations, and the corresponding actuators are activated simultaneously. Moreover, to identify the dynamic behavior of the whole, cracked system, an energy function with different weighting coefficients is introduced to propose a suitable criterion for the performance evaluation of the second approach. Simulations for different crack numbers and locations and external disturbances as a comparative study (in MATLAB/Simulink) demonstrate suitable criteria to determine the number and locations of actuators and reduce the power costs in modern spacecraft in high-precision missions.

Changes in altitude from the sea level have a significant effect on the performance of the internal combustion engine. Turbochargers can be used to maintain engine power by changing the height. In order to test the combination of turbochargers under the conditions of high altitude, the pressure and temperature should be created, which is possible by placing all the test set in the altitude simulator chamber, or at least the conditions of inlet and outlet must be controlled and the pressure and temperature associated with it altitude made. Creating these conditions is difficult to control. The purpose of this paper is to provide a simpler and more cost-effective way to create a r turbocharger test bed on an internal combustion engine in altitude conditions. In this paper, one-dimensional simulation engine is engineered and compared with the motor test results. Using a compression ratio, mass flow rate and corrected distances obtained from simulation, a simple method for testing the required pressure ratio at the desired height is presented. This preliminary design can specify the scope of turbocharger control overhead and reduce the cost and time of operational testing of the bird's device for extracting control overflow.The simulation of the target engine in the research and comparing it with the experimental results at different periods shows a maximum simulation error of 10%. This research is the applicable over a wide range of altitudes. The turbocharged motor is capable of maintaining 90% power up to a height of more than 12.2 km.

In the present study, the spraying behavior of diesel and kerosene fuel in a cylindrical fixed volume combustion chamber for a injector orifice has been investigated. In order to investigate the effect of injector geometry on fuel spray characteristics and atomization quality, different geometries have been used. For this purpose, the microscopic and macroscopic properties of the diesel and kerosene fuel spray for different geometries of orifice compression ignition injectors are modeled and investigated using AVL-Fire. Firstly, the liquid fuel flow inside the injector with cylindrical and converged conical nozzle holes have been modeled and then in the following diesel and kerosene fuels have been used in the grooved nozzle hole. Numerical results show that in this case, the kerosene spray has smaller penetration length and bigger cone angle than diesel fuel. Controlling the properties of the fuel spray is important to increase the combustion efficiency of the engine and also to reduce their pollution, and can be done by changing the geometry of the fuel injector nozzle.

In this paper, first the turbofan engine is modeled and then the fuel flow controller is designed based on the Min-Max algorithm. The engine control system is designed as a dual channel, which means two controllers independent of each other but connected to each other. In the Model-in-the-Loop (MIL) test, the controller and the engine model are run on the computer, but for the Hardware-in-the-Loop (HIL) test, the controller is implemented on an Arduino board that connects to the computer model via USB cable. The main purpose of the controller is to comply with engine restrictions, to provide fuel flow based on the pilot's desired trust in the minimum time and no rapid changes in fuel flow. The results of MIL and HIL tests were evaluated for different inputs, which shows that the controller correctly observed the above notes. The only difference between the HIL and MIL test results is the lower speed of the controller in the HIL test.

Plasma actuators application in flow control was a focus of attention in recent decade. Dielectric Barrier Discharge (DBD) plasma actuators are a subgroup of plasma actuators owing simple structure and performance, fast response time, low power consumption, Low Stealth and lack of moving parts. In this study, the effects of using multiple DBD plasma actuators at different distances and numbers to produce aerodynamic forces were experimentally evaluated on a flat plate in air atmosphere. DBD plasma actuator performance was directly related to the voltage and frequency of the electric current; by raising the voltage above the breakdown voltage, the rate of thrust and lift forces will increase; while the frequency of electric current significantly influences the breakdown voltage. Furthermore, the study reveals that the distance between the DBD plasma actuators has a major role on the amount of aerodynamic forces, offering more impact in comparison with number of DBD plasma actuators.

The study was done to identify periodic Lyapunov orbits in the presence of the primaries oblateness applying Poincaré map at the restricted three-body. Governing equations of the PRTBP orbital motion were derived using principles of the Lagrangian mechanics. Since the governing equations have no closed-form solution, so the numerical method must be applied. So the problem can have different periodic or non-periodic responses to the initial conditions. The proper initial conditions were obtained from combining the third-order approximation of the Unperturbed Restricted Three-Body Problem and the orbital correction algorithm in previous researches. This method required complex and time-consuming mathematical calculations. Therefore, in this paper, the suitable initial conditions of periodic Lyapunov orbit are suggested to identify with the Poincaré map. Poincaré maps are a valuable tool for capturing the dynamical structures of a system, such as periodic solutions via a discrete and lower-dimensional representation of the dynamical flow. The center and boundaries of the islands created in this map are considered as suitable initial conditions to meet the periodic responses. To validate the proposed method, the perturbed Lyapunov orbits family is plotted. Also, in order to illustrate the effect of perturbations, the initial conditions of the perturbed and unperturbed models are compared due to the same initial guess vectors.

This article introduces PCO formation. The PCO array is extended to three satellites, and one satellite is assumed to be the leader and two satellites to be the follower. A new controller has been introduced to control the position (distance) of the follower satellites in relation to the leader satellite and to reject environmental disturbances (atmospheric drag, gravitational perturbations, solar radiation pressure, etc.). The proposed controller is the Adaptive Sliding Mode Controller (ASMC) and somehow based on feedback linearization and the PI controller with adaptive coefficients for both stabilization and tracking. This controller is nonlinear and its stability is proved by Lyapunov method. To evaluate the performance of the proposed controller, first the method of linearization of nonlinear dynamic equations of position of satellites is presented then LQR method is used in the simulation for these equations and the results are compared with the proposed controller. The simulation results show that the proposed controller performed well in the initial formation of the PCO arrangement and also no collision occurs due to the desire relative positions of the follower satellites. The proposed controller performs better than the LQR method in keeping the PCO formation and reject environmental disturbances.

In this paper, the finite horizon orbital optimization of a satellite launch vehicle based on nonlinear predictive model is presented. The proposed method optimizes the flight path variables such as the angle of attack parameters as well as the optimization of the trust value of each stage to achieve the maximum orbital altitude. In this method, the nominal path of the space carrier is obtained and compared with different optimization methods such as pattern search, sequential square programming and genetic algorithm. The resulting path is obtained based on the nominal conditions of the space carrier and by optimizing the variables of the angle of attack function. In this method, using the finite horizon optimization method, the optimal trust in each stage is accomplished with assuming that the total specific impact of each stage is constant. The flexibility of this method in solving optimization problems and the possibility of considering different flight path constraints are some of the advantages of this proposed limited horizon method. The proposed algorithm is used to optimize the orbital height of a native space carrier, and the simulation results show a 24 km increase in its orbital height relative to its nominal path conditions.

The geometric method is one of the nonlinear methods of large-scale relative motion that is used for all the eccentricity and relative distances. In this paper, the relative motion equations for perturbed orbits are developed in the presence of third body. The purpose of this paper is to design a control law to track and control the relative attitude of one satellite in the presence of position dynamics and uncertainty of the moment of inertia. For the orientation of satellite, it is first necessary through the theory of relative motion, to obtain some relative parameters such as position and relative velocity, and so on. The geometric method is used to obtain relative parameters. After obtaining relative parameters, we must track and control the dynamic equations for the target trajectory. Due to the uncertainty in the dynamics of the system, a robust controller must be used to obtain control law. Sliding mode control theory is used to obtain control law. The uncertainty in the satellite inertia is considered due to fuel sloshing during this maneuver. Finally, an appropriate control law is designed that is resistant to uncertainties and has good accuracy.

Nowadays, aided Inertial Navigation System (INS) is implemented to increase accuracy of navigation. However, in some cases, it is not possible to use integrated methods, or if the link to the navigational aiding system is lost, INS should be able to operate standalone with the required accuracy, until the link is accessible again. Therefore, the accuracy of INS should be enhanced to the highest extent. Some of the most significant factors in reducing the accuracy of INS are: measurement noise, inertial sensors error, misalignment, and sensor installation error on the body of vehicle. In this article, methods of compensating these factors are presented to increase the accuracy of strapdown inertial navigation systems consisting of MEMS sensors. First, the effect of noise cancelation on accuracy of INS has been studied in practice. Then, a method is presented for calibration of MEMS accelerometers and gyroscopes. In addition, compensation of installation error and sensor misalignment with body-frame coordinate is studied. Efficiency of the compensating methods has been evaluated and reported in practical implementations.

Much studied have addressed Path planning as one of the main topics in unmanned aerial vehicles; which has yielded different results due to the existing conditions and limitations. To this end, the present study proposed an efficient algorithm underpinned by the propeller optimization algorithm, which had a utility function and could optimize multiple responses simultaneously. BOA is different from other meta-heuristic algorithms as each propeller produces its own unique fit in the path by combining information extracted from different sensory receptors. Accordingly, BOA can solve multi-objective problems. A 3D objective function was used in this study to estimate the length of the shortest path and the intensity of collisions with obstacles, avoid collisions and enhance UAVs’ operational capacity as a function of consumed energy. Furthermore, A smart launcher factor was also included in this algorithm to prevent trapping in local optimizations and promote network coverage in the routing process at the same time. The launcher prevents collision with obstacles in UAVs by adopting geometric techniques and the contour line. The performance of the proposed algorithm was compared with that of the most practical meta-heuristic algorithms (namely ACO and PSO methods). The findings revealed that the BOA algorithm compared to the other two algorithms had the lowest cost and the second lowest cost under the best and worst conditions, respectively. The findings also confirmed the better performance of BOA than the other two algorithms regarding execution time and the optimal value of the fit function.

This article deals with dynamic modeling and nonlinear optimal control of an aerial vehicle taken into account nonlinear modeling of its thruster engine. To do this, nonlinear flight dynamic equations of the aerial vehicle are derived considering nonlinear equations of the thruster engine. A relation between the thrusting force and the fuel consumption of the air-breathing engine is stated and coupled with the nonlinear dynamic equations of the vehicle. Presenting the formulation of the state-dependent nonlinear optimal control, the nonlinear dynamic relations of the aerial vehicle are considered as constraint equations of the optimal control and a cost function including state and input variables is defined. Then, the Hamiltonian function of the optimal control problem is formed and optimality equations obtained. By numerical solving of the problem, various parameters like as the path angle, uncertainties, attack angle and weighting coefficients of the optimal control are considered and several simulations are presented. The results show that increasing the attack angle leads to increasing of the control time and effort of the system. Also, changing the weighting coefficients lead to various optimal paths as increasing the input weighting coefficient decreases the fuel consumption. It is seen by changing the uncertainty parameter of the model, transient response is changed but finally the optimal control method is able to track the desired path. Furthermore, by changing initial conditions and parameters, the nonlinear optimal control of the system is effectively performed which indicates the priority of the proposed method