جستجوی مقالات مرتبط با کلیدواژه « ماهواره » در نشریات گروه « مکانیک »

Due to the limited communication time between LEO satellites and the ground station and large volume of satellite image data, it is necessary to establish a high transmission rate connection. Therefore, the use of a high-efficiency satellite antenna is inevitable. In order to use the limited viewing time of the satellite, it must be possible to communicate with the ground station at low angles and change the main beam direction of satellite antenna to the direction of the ground station. The proposed array is composed of 8 antenna elements with a compact microstrip feed network. Low profile structure, circular polarization, high gain and low cost specifications of this proposed antenna array, candidate it for conformal array antenna in satellite applications.

Components of the LEO environment consisting of high vacuum, ultraviolet radiation, thermal cycles, Atomic Oxygen, charged particles, electromagnetic radiation, micrometeorites, and space debris significantly degrade the material properties of polymer matrix composite materials. Therefore, in the design of space structures using composite materials, the possible deterioration of mechanical properties due to long-term exposure in LEO space environment should be carefully considered. Accordingly, a reliable understanding of the LEO environment in its effects on composite materials as well as investigating its properties is very important. The space simulator system under vacuum is designed with the aim of creating a standard environment with the ability to reach high vacuums and also the ability to create low temperatures, as well as to check the qualification of space components and research materials used in the satellite. These systems analyze the behavior of satellite components, evaluate their thermal balance and performance to ensure mission success and survivability.

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.

The main design goal of thermal control subsystem is to keep the temperature of the satellite equipment within specified margins. Satellite experienced extreme hot and cold conditions during the mission, so it is difficult to choose the materials and surface coatings for these conditions. If design is done to deal with hot conditions, satellite power for keeping warm in cold conditions is increased and this is while supply of required power for heating the satellite in cold conditions is one of the important design concerns. Due to restrictions in space systems for power supply, the value of power for thermal control subsystem should be optimal. This paper describes the process of designing a thermal control subsystem of small satellites using a simple model. In the present study, the thermal design is done by using the optimal choice for the radiator cover. This design has been done in a way that the least amount of heater power is required. For this purpose, optimization and linear programming have been used. Simulation results show that satellite temperature in warm and cold conditions is in the range and this method significantly reduces the amount of required power heater.

According to the technical specifications of the future generations of telecommunication (the fifth generation and later), which should provide new services with very high data rates in the minimum time and a wide coverage, as well as the exponential increase in traffic, the use of combined space-air networks Land is essential. It should be noted that the management of this type of combined networks has major challenges in providing such services. Meanwhile, the intelligent management of resources in satellite-based hybrid networks will lead to increased capacity and improved service quality. For this purpose, in this article, a comprehensive review of the use of artificial intelligence in the field of satellite communications will be discussed. In the field of intelligent increase of capacity, various factors such as how to configure the network, how to allocate resources such as spectrum, energy and power will be investigated with consideration of intelligent interference management. Finally, in the field of service quality improvement, factors such as how to model and intelligently predict traffic, as well as how to deal with harmful environmental conditions, will be presented.

In this paper, a step-by-step laboratory procedure for performing a satellite's payload’s alignment measurement is presented. Four highly accurate theodolites are used along with two or more alignment corner cube to accurately extract the final attitude. Theodolites are arranged around the satellite in such a way that they have a clear direct view of the alignment cubes mounted on the payload and the satellite. Two theodolites should point to the payload’s alignment cube and the other two theodolites must point to the satellite’s alignment cube. Each theodolite must see at least one other theodolite, directly. Finally, by forming the coordinates systems of the payload and satellite in the theodolites coordinate system along with using the coordinate transfer matrices, the payload alignment correction matrix will be extracted in detail. The total method accuracy is within the order of few arcseconds.

In this paper, Design, optimization, and implementation of a Transparent Pseudo-Noise satellite Ranging relay are presented. CCSDS and DSN standards were studied for requirements of on board hardware for LEO satellites and ground station. To optimize the performance of the relay, a ground station for a two-way ranging system was modeled in MATLAB according to the DSN standard architecture. The effects of non-ideal hardware for the Implementation of an on-board relay were simulated. The simulation results confirmed that the designed transparent relay hardware satisfies the requirements of CCSDS and DSN  ranging standards system and has no detrimental effect on it. Finally, the implemented relay was tested with a Transceiver similar to the ranging Transceiver was confirmed for operation.

The satellite system is subjected to shock during launch and separation. The shock causes damage to various parts, including the electronic system. In this article, the electronic board of the satellite guidance system has been analytically examined. This board is located on 5 legs that are connected to the body. The purpose of this study is to design the legs so that the least amount of shock is introduced to the electronic board. The electronic board is assumed to be a plate. Effective parameters in shock transmission, including geometrical parameters and the leg angle with the horizon surface, have been studied. The influence of the leg angle with the horizontal surface on the stiffness of the system has been investigated using analytical relations. Finite element simulation has been used to validate the equivalent stiffness. The results of this study show that changing the base angle causes the transmissibility of acceleration to change according to the excitation frequency. Acceleration transmissibility depends on the ratio ω/ω_n . For this reason, different parameters should be selected according to the excitation frequency range. If different parameters are not selected properly, the output acceleration will be higher than the input acceleration.

The interaction of thruster plumes with satellite components can have undesirable effects, such as disturbance force/torque, thermal loading, and species deposition in the surfaces. The purpose of this paper is to use the Direct Simulation Monte Carlo (DSMC) method to analyze the 3D plume impingement flows and investigate its effects. Two impingement problems are computed. The impact of a jet of nitrogen on an inclined flat plate is considered. Good agreement is found between surface quantities calculated by DSMC and experimental data. The plume of a hydrazine control thruster firing in a model satellite configuration is simulated. Surface quantities and net impingement effects are calculated. The effects of partial displacement of the thruster locations on the results have also been investigated. The results show that a 20% displacement of the thruster location can change the disturbance force/torque by up to 15% of the initial values.

Man's entry into outer space and specifically the use of satellites does not have much history compared to airplanes and ships. This has caused many uses of satellites in space to be associated with various legal uncertainties. The governments with satellite technology claim to use satellites in order to strengthen their interests and national security and based on this, they have prioritized military and weapons satellites; , they want the peaceful use of space for satellites and the prohibition of weaponization of this space. The main question in the present article, which is done in a descriptive-analytical way, is that in the legal conflict of peaceful use of satellites by weaponizing outer space and emphasizing security components, which one should prevail? The result of the current research is that the principle of peaceful use of satellites in the mentioned space has gradually decreased in importance and due to this, it is no longer possible to consider peacefulness as a basic security-oriented component in the field of using satellites. and this allows other governments, including Iran, to take advantage of their security interests in outer space by complying with their requirements according to the United Nations Charter and in order to legitimately defend their security interests. It seems that as a suggestion, it is possible to refer to the necessity of COPUS, which should prepare and approve a binding document regarding countering the weaponization of space through the use of military satellites. governments to deliver

The use of accurate and reliable navigation systems for determine the exact position and attitude of aerospace equipment is of great importance. Magnetometer is one of these types of sensors which is a low cost and reliable tool for determining exact position. The purpose of this study is based on use a magnetic field sensor to determine the position and estimate the orbit of a typical satellite. By using magnetometer measurements and an independent algorithm, the position of the device is first estimated. This is done by reading magnetometers and using models of the Earth's magnetic field. The use of the above technique is associated with errors that have been reduced by using the extended Kalman filter. The filter used overcomes the initial errors and provides the desired accuracy approximation for determining exact position of device. The results for a circular orbit will indicate the high accuracy of these estimates for low-altitude orbiting satellites.

The thermal control subsystem (TCS) of satellites is responsible for providing and maintaining the allowable temperature of that satellite equipment. Naturally, if this subsystem is not properly designed, other subsystems, including the payload, will not be able to do their duties. In general, there are two active and inactive methods for designing TCS. Studies show that the elements can be suitable for controlling the temperature of a small-scale satellite (SSS) which has low weight, volume, and power consumption. Therefore, passive temperature control this equipment is widely used in SSS. The most common of these equipment are coatings, insulators, and paints. But some active elements are commonly used to make the thermal design of small satellites safer. Heaters are more common among thermally active elements, because they .have less volume and weight than other active equipment, and consume less power than other elements

In order to reduce the weight and increase the efficiency of thermal control subsystem of the small satellites, employing efficient equipment is necessary. Smart radiator is one of the useful equipment in this field. Emission coefficient of these radiators change as a function of temperature or voltage. These variations cause heaters consumption power to be decreased in satellites. In this study, smart radiator is simulated and used in a thermal control subsystem of a small satellite. The result is compared to a conventional radiator. Non-linear proportional-integral method is employed to control the temperature. In this method, each part temperature is considered as state variable and heaters are considered as operators. Simulation results show that non-linear proportional-integral control method has better performance in decreasing the overshoot and settling time in comparison with the linear proportional-integral control method. Besides, smart radiators cause heaters consumption power to be declined. In a way that for the battery about 7% and for the top plate about 27% of the heater power consumption has been reduced.

The heat control subsystem of each satellite is responsible for providing the allowable temperature of the equipment of that satellite. Naturally, if this subsystem of the satellite is not designed properly, the satellite mission will not be performed properly. In the present work, small-scale satellite heat control subsystems have been designed by using Thermal desktop and Sinda softwares. This simulation is based on two cold and hot critical states and the results are compared with the available experimental data in the literature, which has about 9% difference. According to studies conducted in this research, the most suitable heat control equipment are paints, coatings, insulators and heaters. According to the analysis, the cargo subsystem experiences the temperature about -19 to 28 degrees Celsius before applying the proposed thermal scheme, which using AMJ-700-IBU paint, its temperature is in the range of -10 to 47 degrees Celsius that is within the allowable temperature range.

In recent years, according to progress of aerospace industries particularly satellites it has been paying much attention to attitude determination of satellite. Attitude determination of satellite in various ways, such as: radio, radar, optical methods and using GPS, q and kalman filter is done, that each of these methods has advantages and disadvantages. Due to disturbance in space, high accuracy in attitude determination, various algorithms to obtain properly accurately attitude, desirable access to position, velocity and time of satellite to achieve desired attitude determination satellite are considered a top priority. However in this Paper, first we review the various methods for attitude determination satellite and then examine of operation and relations of algorithms discussed. This Paper focus on the advantages and disadvantages of qEKF algorithm in compare to other methods are available.

In this paper, an accurate computational method is proposed to measure the relative accuracy of the star sensor, which does not require complex laboratory instruments. The proposed method uses the direct observation of the night sky along with the exact equations of motions of the Earth and stars to measure the accuracy of the star sensors in the order of 1 arcsecond. The Classical Laboratory Measurement Procedures of the star sensor’s accuracy requires at least a sky simulator along with some auxiliary tools such as a collimator, an accurate 3 DOF Rotary table, an exact alignment instruments, and so on. The classical procedure, in addition of being complex and time-consuming, is costly, and the auxiliary tools also increase the measurement error by themselves. Identifying and eliminating these errors are more difficult process. The proposed procedure is more accurate and more reliable because the sensor is tested in its actual operating environment, i.e., the sky, rather than the simulated laboratory environment.

This paper is focused on the application of transparent antennas to address the issues related to the unpleasant appearance and non-aerodynamic behavior of traditional antennas. Transparent antennas are divided in two main categories. In the first category, the resonant structure and the ground plane of an antenna, which is usually made of metal, is made of thin film transparent materials or a metallic mesh. In the second category, the structure of a dielectric resonator antenna (DRA), which is made opaque dielectrics, is realized, using a transparent dielectric, such as glass. In this paper, common types of transparent antennas and the methods of realizing each of these antennas are reviewed. The paper also discusses different potential applications of such antennas, especially space applications to save the limited available mass budget in small satellites.

The CANSAT design, a simple and small scale of a satellite, is an experience for preparing for the design of a Life cycle of the space project. In this paper, the process of design and construction of ARTA CanSat, which is participated in the scientific-exploratory class of international competition in CANSAT Iran, has been written. The operation scenario is that an automatic scanner system, lands after releasing from 300 meters above ground level with using a recovery subsystem (parachute), which is controlled its downfall and landing position by the parachute controller system. During the descent, the data is transmitted by the sensors and sent to the ground station. After touchdown, the Hotwire system's operates and separates the parachute from the Rover section; Then CanSat with the moving on the ground by using the simpler section embedded underneath the structure, performing excavation while moving to the target point. Throughout the mission, the health of the biological payload is preserved.