چكيده به لاتين
Nowadays the use of many modern technologies are not possible without space. On the other hand, human curiosity to know more about the universe and access to other planets to study them, doubles the importance of this issue. To achieve this goal, a space mission should be defined and designed. One of the most important steps in this process is the satellite trajectory deisgn. The design of the trajectory of the satellites depends on the type of their propulsion system. The propulsion systems in satellites are considered as impulsive and low-thrust propulsion system. To this reason this dissertation consists of two parts. In the first section, it is dealt with to the trajectory design of the on-orbit servicing satellites with impulsive propulsion system. Studies reveal that in the planning of on-orbit servicing missions, the orbital elements of the parking orbits, including the eccentricity, the inclination angle, the right ascension of ascending node, the argument of perigee and the phase difference between servicing satellites on the parking orbit, are determined by disgner, while these values may not be adequate to achieve the minimum fuel consumption. For this reason, in this dissertation in order to improve and supplement the existing methods, it is assumed that in addition to the parameters "the time the servicing satellite leaves the parking orbit", "duration of orbital transfer", "duration of phasing maneuvers of each servicing satellite", "the time required to perform the assigned task on the target" and "the position of the servicing satellites on it", are determined through the optimization process. Also, it is considered a special type of non-cooperative rendezvous maneuver for meeting the servicing satellite and the target in the final orbit, which makes easier to perform servicing mission. Since the mission planning is intended to minimize fuel consumption, a particle swarm optimization algorithm has been employed to achieve this goal. Applying the proposed approach resulted in a reduction of about 20% in the total velocity changes required to execute the on-orbit servicing missions studied in this dissertation. In the second part of the dissertation, it is dealt with the design of satellite with low-thrust propulsion system. In general, the methods that exist to solve the optimal trajectory design problem of satellite with low-thrust propulsion system are divided into two groups of direct and indirect methods. One of the problems in employing these methods is that they are highly dependent to the initial guess. On the other hand, during the preliminary design phase of the mission, a large number of possible trajectories should be investigated; therefore, the search space of the problem will be very large and it is very time consuming to pass through this stage of mission design, through these methods. For this reason, in this dissertation, an approximate and near-optimal solution is presented, which can be used both as an initial guess of direct and indirect methods and as a very effective method for the preliminary design phase. The main idea of this approach is to consider the thrust angle as a Fourier series with limited and uknown coefficients obtained through an optimization algorithm with the aim of the minimizing mission duration. This method has been used to solve the trajectory design problem in the two-body and the three-body problems. The results show that the proposed method determines the number of revolutions and the time of the mission with high accuracy. It provides the number of revolutions without error and the mission time and fuel consumption with an accuracy of less than 2% compared to indirect methods.
