نشریه علوم، فناوری و کاربردهای فضایی
پیاپی 1 (بهار و تابستان 1400)

In all types of satellites, communication systems are utilized for data transmission between satellite and ground stations. pointing the communication antennas to the ground is necessary for the correct mission transmission information. The vibration of the satellite antenna leads to the deform antenna pattern, creating noise and reduction of connection quality. Moreover, working the attitude control actuators near the antenna's natural frequency leads to its resonance and large amplitude vibration in the antenna and satellite structure. Thus it is necessary to identify the satellite antenna dynamic behavior as natural frequency and damping ratio. In this paper the satellite antenna is intended as a smart beam, based on the free vibration of clamped-free beam shape of satellite antenna and sensing its vibration by the piezoelectric sensor, its dynamic characteristic as damping and frequency is identified and verified by comparing the results with experimental ones. The considered mathematical model has very accurate and this model can be used to determine the dynamic behavior of the antenna in different satellite secondary structures.

In this paper, we investigate the effect of task scheduling on the lifetime of a real-time hard drive system that uses a composite energy tank consisting of a battery, a supercapacitor, and a solar energy picker to power supply itself. The life of a system is defined in terms of the moment the system starts up to the moment its tasks are disrupted due to lack of energy. Due to the nonlinear properties of batteries and supercapacitors, which cause their internal charge to be divided into available (IAC) and inaccessible (IUC), the lifetime of such a system depends entirely on the charging and discharging pattern of the energy tank, as this pattern ultimately leads to the amount of charge stored in the IUC section and the amount of charge extracted from this section. Therefore, we can influence the lifetime of the system and increase it by managing the charge / discharge pattern of energy tanks. Since the pattern of energy delivery from the environment is beyond the control of the system, we try to influence the tank charge / discharge pattern by adjusting the consumption pattern to ultimately improve the life of the system. In this regard, we have presented two scheduling algorithms MCF and MGF, which are respectively trying to perform the most consumed and least consumed task in the system, and then using the MCG policy, which at any time according to the conditions for using one of these decision algorithms are presented. Experimental results show that we can increase system lifetime by between 5% and 16%. Considering that in recent years the issue of using supercapacitors along with batteries and solar cells in space systems has been raised, so the results of this research can be investigated for use in satellites.

Although regarding low density and high ratio of strength to weight of titanium, the application of this metal is restricted in space industry due to its low surface properties including low wear resistance which could lead to the cold weld. For improving this property of titanium, applying protective coatings are effective. Physical parameters, wear resistance, adhesion of applied layer, and surface microhardness are impressive and could avoid occurrence of the cold weld. Therefore, by applying plasma nitriding on chromium layer in titanium used in space constructions, tribological properties (wear resistance, adhesion and surface microhardness) were studied. Results confirmed the formation of nitride chromium phase on titanium substrate due to the applying plasma nitriding on chromium layer and increasing the surface microhardness up to 1109 HV and improving the wear resistance and adhesion subsequently. The value of the friction coefficient was decreased down 0.16 which could avoid the cold weld occurrence.

In complicated systems such as satellites, each subsystem's design can affect the design of the overall system. In the design procedure, the effect of each technology on the other components should be determined. Because of the increase in space trash, the propulsion subsystems are introduced to avoid collision in space. One of the methods to attain high altitude orbits is flight maneuvers using propulsion systems. Several types of propulsion systems are utilized in satellites. Resistojets can be employed as low-cost propulsion systems for satellites because they do not use complicated technologies. In the present research, a resistojet propulsion system is designed for a CubeSat for the mission of orbital altitude reduction. The propellant is selected considering all properties. The design of the nozzle and the heater is also introduced. The overall layout of the system is presented. In the end, an algorithm is proposed for electrical propulsion systems for a specified mission.

Various sensors can be used to attitude determination of a satellite, including the Earth Horizon Sensor. These sensors generally divided into two static and scanning types. In the static type, a two-dimensional array or several linear arrays capture instantly the Earth image information to attitude determination. In the scanning type sensors, a narrow-band optical beam mechanically sweeps a ring of space to find the crossing point of the beam to the Earth using a single spot detector. In this research, a new algorithm for a static earth horizon sensor with a two-dimensional array is presented. The algorithm does not need elevation of the sensor and this is the main advantage with respect to old ones. To extract the equations, a tangent vectors from the aperture of the pine-hole camera to the Earth is considered and this vector is extracted in two different body coordinates and base coordinates that are connected to the ground. By equating the two, an equation with three unknowns of pitch, roll and elevation is obtained. The equation is satisfied by each pixel on the Earth horizon Thus; the system of equations is formed by the number of the Earth horizon pixels in the camera image and by numerical solving of the equations unknown parameters are estimated.

One of the most critical design points in satellites is to achieve the minimum mass (or weight) by meeting all the requirements and constraints (constraints on strength, location and pointing vibrations). In this regard, the structure can play a very important role because the structural designer has more freedom of action in determining the design of the structure compared to other subsystems. In the present study, the process of designing the structure of a small satellite, with the aim of achieving the lowest possible weight and maintaining the prevailing requirements (vibration constraint), has been investigated. The process used to achieve the above goal is a change in the geometric dimensions of the structure. Satellite modeling steps with all subsystems have been performed in solidwork software and in ANSYS software, modal analysis, strength and random vibrations have been performed. Also, using the optimization of genetic algorithm in this software, the geometric parameters of the structure such as the thickness of the frame have been obtained in such a way that the structure has reached the lowest possible weight and by satisfying the prevailing conditions for vibration constraints, strength and random vibrations. The results show that by choosing the right thickness of frame, the weight and frequency of the first satellite mode can be significantly reduced, and at the same time, the optimal final structure satisfies all the constraints applied by the launcher and has sufficient strength and rigidity. Is.

Ti-6Al-4V is one of the most common materials in the aerospace industry. For example, satellite fuel tanks are made of this alloy. Among manufacturing processes, forming processes is one of the most widely used areas in the manufacture of Ti-6Al-4V components. Due to the importance of determining the allowable deformation limit in the successful design of the Ti-6Al-4V forming process, in this paper, the amount of critical damage was studied. For this purpose, parts with double-cone geometry with grooves on the maximum diameter with two initial microstructure, lamellar and equiaxed, were fabricated and subjected to hot compression testing. The results showed that the initial equiaxed microstructure provides good accumulated damage tolerance. Up to 2.38, 2.67, and 5.89 accumulated damage values, according to Cockcroft-Latham, Brozo, and McClintock criteria, respectively, no crack was observed on the samples. However, with an initial lamellar microstructure, the damage tolerance was significantly reduced. The critical damage value based on Cockcroft-Latham, Brozzo and McClintock criteria was 1.05±0.02, 1.03±0.02, and 2.56±0.05, respectively.

In this paper, a task scheduling and mapping method based on multi-objective particle swarm optimization (MOPSO) algorithm is presented to improve lifetime reliability of multiprocessors systems on chip. This method considers power consumption temperature and performance along with the lifetime reliability due to the antagonistic relations of these parameters. These antagonistic and dependent relations make the design space exploration and optimization processes complex. The proposed method is based on MOPSO algorithm and starts with an initial population of candidate solutions which are generated randomly and represents valid task scheduling and mapping solutions. By changing the scheduling and mapping parameters during the MOPSO algorithm, new solutions are produced and the design space is explored based on the objective of the target problem of this method. Several experiments on random and real-life benchmarks are performed to verify the effectiveness of our proposed method. The results demonstrate the capability of the proposed method in effective design space exploration and generating the Pareto front. Moreover, comparisons to the related research show 35%, 23%, 19% and 3% improvements in performance, lifetime reliability, temperature, and power consumption on average.

The microwave electrothermal thruster is the propellant that converts the microwave energy into thermal energy. Nowadays, with the development of technology, electric thrusters are very important in terms of producing high specific impact and low consumption. These thrusters are capable of producing acceptable thrust over a long period of time that are suitable for helical orbital transfer missions. In this type of propulsion, the propellant gas is heated and expanded, which produces propulsion force. In this paper, the conceptual design process of a microwave electrothermal thruster is presented. The propulsion system includes propulant gas, propulsion storage tank, amplification chamber and power generation source, which includes batteries and solar arrays. In the text of the article, the method of calculating the mass and the specifications of each part are presented in detail. Finally, in order to validate the conceptual design process presented in this study, the necessary studies have been discussed.Conceptual design has been done for a 100 kg satellite, which is desirable to travel in a week from an orbital height of 300 to 800 km during a spiral treansfer. The propulsion system and mass of each subsystem are obtained.

Space systems in Low Earth Orbit (LEO) are exposed to the destructive parameter of atomic oxygen. In long-term missions, the rate of degradation of the material resulting from the reaction with atomic oxygen is significant and reduces the performance of the structure. Due to the harmful effects of atomic oxygen on materials, the choice of atomic oxygen resistant materials or the use of durable surface coatings is very common. In this study, the corrosion resistance of atomic oxygen of an interconnector part of a solar cell was studied by applying a silicone base coating. In order to investigate the corrosion behavior of atomic oxygen, ground test method with equivalent conditions of LEO orbit was used by DC plasma equipment. Initially, the parameters of the atomic oxygen corrosion ground test were determined under the equivalent conditions of the LEO orbit. The results of atomic oxygen application in this study showed that the amount of atomic oxygen erosion yield of silicon coating is significantly lower than the amount of atomic oxygen erosion yield of silver substrate. Also, the study of the coating surface after applying atomic oxygen by SEM images led to the determination of the optimal coating thickness. EDX results showed that after applying atomic oxygen, no significant change in the chemical composition of the coating was achieved.