In the mission of satellites, instantaneous positioning and estimation of the future position are necessary. In the communication satellite, this matter is so important. Thus in the Tadbir satellite, for the first time in the country, we consider an Orbit Determination subsystem that equipped with a space borne GPS receiver and contains appropriate algorithms in order to achieve satellite position data during the lunch period and in the Orbit. In this paper a brief review of the software and hardware parts of this subsystem is presented. In additional the process of testing to achieve good performance, including functional tests of ODS processor board and functional test of GPS receiver with GPS simulator, environmental condition tests, mechanical tests, thermal vacuum cycle tests, electromagnetic compatibility test and finally integrated satellite tests are stated. This paper, in addition to a description of the Tadbir satellite Orbit Determination subsystem, an implementation testing of a satellite subsystem is demonstrated, which can be useful to other space field researchers.

In this paper, Simultaneously Attitude and Orbit Determination of satellites using only star tracker measurement is investigated. At first, the 6 D. O. F. governing equation of the system considering the gravity force and gravity gradient torque is obtained using the Newton-Euler equations. Next using a simplified star catalog, star tracker output in the desired Orbit is calculated. To estimate variables of the Attitude-Orbital equations, an estimation algorithm based on the UKF is proposed. In each step using the governing equation, estimated parameter and star tracker output, satellite angular velocity and orientation of body against NED frame are estimated and these calculated parameters are used as input for the next step. According to Results after about one revolution of Orbit, the estimation is converging and the ratio of filter’ s solution error to dynamic’ s solution error for the roll and yaw angles is about 2 percent and for the pitch angle and angular velocities are about 33 percent. In order to reduce position and linear velocity errors, the filter needs more time and at the end of 6 cycles, the performance shows 40 percent improvement in average.

This paper is dedicated to determining the Orbit of a satellite by using Unscented Kalman Filter (UKF), in which, a GNSS is used as the observation sensor. During this goal, firstly we have simulated the satellite Orbit; considering oblations effect. Secondly, exploiting a nonlinear model of Orbit dynamics, preliminary Orbit Determination is prepared via UKF algorithm. The range between the satellites and space vehicle as the parameter of observation in the filter is obtained. Afterwards, the estimated preliminary output data is corrected and more precise position of the GNSS satellite is determined based on predicted observations errors. Consequently, the outcomes of the research exhibits the acceptable satellite Orbit Determination error range.

In recent years, the development in the space industry and the ability of building, launching and infusion of satellites in the lower Orbit has put the limited number of countries with such technology. In order to complete the entire cycle of the space industry, the satellite navigation and control, which have been neglected since the beginning of the movement of space science, has to be considered specially. The Orbit Determination in one sentence is the application of a variety of techniques for estimating the Orbits of objects such as the moon, planets, and spacecraft. In dynamic astronomy, the Orbit Determination is the process of determining Orbital parameters with observations. In particular, Orbit Determination of planets of solar system is adjustment of noisy Orbital observation that consist of random and systematic error for force models and estimation of model parameters by observation (In order to access a mathematical model that illustrates the path of the celestial object in the path before and after the observation time). To simplify, this process is divided into two parts. First, the initial Orbit is estimated and then make corrections to the determined Orbit. The purpose of initial Orbit Determination of object that is moving around earth, is calculation of object Orbital parameters by a few observations; furthermore initial Orbit Determination is used for detecting missing object in space. To determine the precise Orbit, it is necessary to determine the initial Orbit with good accuracy, which indicates the importance of the initial Orbit Determination. Different type of observations is used to make an initial Orbit Determination in which observations can be collected by ground stations that contain angular angles, elevations, distance, and distance rate. These observations are made by the radar and the telescope, because the collection of observations without instrument and naked eye does not have enough precision and sensitivity for Determination of the space object Orbit, but since the extraction of distance observation is expensive and sometimes impossible, angular observation is used. In this paper, a new method has been presented for extracting angular viewing through an optical imaging system. This method is an automatic and efficient method with the ability of real-time data analysis and the base of that is astronomical imaging by CCDs (chargecoupled device). The images captured by this method have a lot of information about stars, galaxy, satellites’ streak, etc. In this paper, automatic method is presented for streak detection which consist of 5 steps: 1) image denoising, 2) extracting of star centers, 3) extracting astronomical coordinates of stars (declination and right ascension), 4) matching between astronomical and pixel coordinate of stars, 5) calculation of satellite streak model. Then, with using the extracted model, the coordinates of beginning and end points are detected. With the celestial coordinates of beginning and end point of streak Azimuth and elevation of satellite on both sides are determined. On the other hand, to evaluate the proposed method and the validity of the input parameters for initial Orbit Determination, the azimuth and elevation values of the beginning and end points of streak can be calculated by precise Orbit file and then these results compare with results of purposed method. Comparing results indicate a difference of about milliseconds.

In this paper, the invasive weed optimization is used to determine the satellite Orbit. The generated TLE can be used in antenna positioning, mission control and monitoring. Also, it can be sent to the satellite for on board applications. In the proposed method, the trial generated TLEs in each iteration are propagated to achieve the satellite position versus the time, and then they are compared with the real data position achieved from the tracking systems to build an error fitness function. Finally, the invasive weed optimization (IWO) algorithm is used to minimize this error. The proposed method is verified by using CHAMP satellite position data, as a real example.

Development of low-cost small satellites has been at the center of attention in recent years. Concurrent Orbit and Attitude Estimation (COAE) requires fewer sensors onboard and subsequently results in some cost reductions. In this regard, the present paper has focused on addressing the importance of COAE utilizing temperature rate on satellite surfaces. To this end, the thermal model for a low Earth Orbiting satellite is introduced first. A three-axis stabilized spacecraft is assumed equipped with small measurement plates that are isolated from each other and from the internal heat sources of the satellite. As the Sun and the Earth are the significant sources of radiation for a near Earth space system, the view factor is the key parameter for observability of the Orbital elements, while the Sun radiation is responsible for the Attitude observability. The Earth albedo factor is a major uncertain parameter required for the thermal analysis of low Earth Orbiting satellites. This parameter is greatly dependent on the Earth’s local terrain and climatic conditions such as instantaneous cloud coverage. To address the problem of albedo factor uncertainty, it is estimated Simultaneously with the Attitude and Orbit of the satellite. NASA's CERES project provides satellite-based observations of the Earth’s radiation budget and clouds over almost 18 years. In this paper, CERES data tables for the Earth’s thermal flux and albedo factor have been used to produce more realistic measurement data. The nonlinear filter of Unscented Kalman Filter (UKF) is also exploited for the state estimation. Lack of sun radiation during the satellite’s eclipse intervals results in the loss of Orbit and Attitude observability. The performance and viability of the proposed COAE algorithm are verified by Monte Carlo simulations. Moreover, a sensitivity analysis is conducted within a wide range of semi major axes, eccentricities, and inclinations. The results demonstrate the high sensitivity of the algorithm to the Orbit altitude and the sun rays direction.

Preliminary Orbit Determination of unknown space object such as satellites that are launched by other countries, military satellites and uncatalogued Space debris is important in space activities. So by means of information of ground based electro-optics systems, Orbital elements of satellite can be estimated accurately. In this paper we studied most famous three methods of classical Orbit Determination analytically which are based on angles-only observations of Orbit. Then, by using observable data and simulations, results of computational codes of three methods are compared and the effect of time separation between observed points of satellite path is demonstrated and results are analyzed.

Comparison of some numerical integration methods of solving the differential equation of motion of a satellite is the main subject of this paper. Since the equation of motion of a satellite is a second order differential equation, therefore, six initial values should be introduced to the numerical solution. These six initial values are the components of position and velocity vectors in an inertial frame respectively. Comparing numerically integrated position and velocity vectors with Keplerian Orbit; one can obtain the bias of the numerical integration method in a satellite-centered coordinate system. In this research, three methods of Runge-Kutta, Runge-Kutta-Nystrom, and the predictor-corrector method of Adams-Bashforth and Adams- Moulton are investigated for a low earth Orbiting satellite. Numerical results show that with integration size of 30 seconds, the Runge-Kutta method, Adams-Bashforth and Adams-Moulton predictor-corrector algorithms, and Runge-Kutta-Nustrom provide closer Orbit to the theoretical Orbit respectively.