In this paper, the satellite in low Earth orbit (LEO) with atmospheric drag perturbation have been studied, where Newton Raphson method to solve Kepler equation for elliptical orbit (i=63 , e = 0.1and 0.5, Ω =30 , ω =100 ) using a new modified model. Equation of motion solved using 4th order Rang Kutta method to determine the position and velocity component which were used to calculate new orbital elements after time step ) for heights (100, 200, 500 km) with (A/m) =0.00566 m2/kg. The results showed that all orbital elements are varies with time, where (a, e, ω, Ω) are increased while (i and M) are decreased its values during 100 rotations.The satellite will fall to earth faster at the lower height and width using big values for ecce

The perturbed equation of motion can be solved by using many numerical methods. Most of these solutions were inaccurate; the fourth order Adams-Bashforth method is a good numerical integration method, which was used in this research to study the variation of orbital elements under atmospheric drag influence.  A satellite in a Low Earth Orbit (LEO), with altitude form perigee = 200 km, was selected during 1300 revolutions (84.23 days) and A_Sat / M_Sat value of 5.1 m²/ 900 kg. The equations of converting state vectors into orbital elements were applied. Also, various orbital elements were evaluated and analyzed. The results showed that, for the semi-major axis, eccentricity and inclination have a secula

     In this research, the eccentricity will be calculated as well as the best height of satellite orbit that can used to transfer from that orbit around the Earth to construct an interplanetary trajectory, for example Mars, when the transfer can be accomplished by a simple impulse, that means the transfer consists of an elliptical orbit from the inner orbit (at a perigee point) to the outer orbit (at apogee point). We will determine Keplerian equation to find the value of a mean anomaly(M) by Rung-Cutta method.

There are several types of satellites orbits around the Earth, but by this study, we find that the best stable orbit to the satellite that is used to inter its orbit around Mars is the Medium Earth Orbit (MEO) at a hei

To transfer a satellite or a spacecraft from a low parking orbit to a geosynchronous  orbit, one of the many transition methods is used. All these methods need to identify some orbital elements of the initial and final orbits as perigee and apogee distances. These methods compete to achieve the transition with minimal consumption of energy, transfer time and mass ratio consumed ), as well as highest accuracy of transition. The ten methods of transition used in this project required designing programs to perform the calculations and comparisons among them.

     The results showed that the evaluation must depend on the initial conditions of the initial orbit and the satellite mechanical exception as well as

An evaluation was achieved by designing a matlab program to solve Kepler’s equation of an elliptical orbit for methods (Newton-Raphson, Danby, Halley and Mikkola). This involves calculating the Eccentric anomaly (E) from mean anomaly (M=0°-360°) for each step and for different values of eccentricities (e=0.1, 0.3, 0.5, 0.7 and 0.9). The results of E were demonstrated that Newton’s- Raphson Danby’s, Halley’s can be used for e between (0-1). Mikkola’s method can be used for e between (0-0.6).The term  that added to Danby’s method to obtain the solution of Kepler’s equation is not influence too much on the value of E. The most appropriate initial Gauss value was also determined to

This research dealt with choosing the best satellite parking orbit and then the transition of the satellite from the low Earth orbit to the geosynchronous orbit (GEO). The aim of this research is to achieve this transition with the highest possible efficiency (lowest possible energy, time, and fuel consumption with highest accuracy) in the case of two different inclination orbits. This requires choosing a suitable primary parking orbit. All of the methods discussed in previous studies are based on two orbits at the same plane, mostly applying the circular orbit as an initial orbit. This transition required the use of the advanced technique of the Hohmann transfer method for the elliptical orbits, as we did in an earlier research, namely

Fuzzy orbit topological space is a new structure very recently given by [1]. This new space is based on the notion of open fuzzy orbit sets. The aim of this paper is to provide applications of open fuzzy orbit sets. We introduce the notions of fuzzy orbit irresolute mappings and fuzzy orbit open (resp. irresolute open) mappings and studied some of their properties. .

In this paper, the Mars orbital elements were calculated. These orbital elements—the major axis, the inclination (i), the longitude of the ascending node (W), the argument of the perigee (w), and the eccentricity (e)—are essential to knowing the size and shape of Mars' orbit. The quick basic program was used to calculate the orbital elements and distance of Mars from the Earth from 25/5/1950 over 10000 days. These were calculated using the empirical formula of Meeus, which depended on the Julian date, which slightly changed for 10000 days; Kepler's equation was solved to find Mars' position and its distance from the Sun. The ecliptic and equatorial coordinates of Mars were calculated. The distance between Mars and the center of the E