Orbital maneuver transfer time is traditionally accomplished using direct numerical sampling to find the mission design with the lowest delta-ʋ requirements. The availability of explicit time series solutions to the Lambert orbit determination problem allows for the total delta-ʋ of a series of orbital maneuvers to be expressed as an algebraic function of only the individual transfer times. Series solution was applied for Hohmann transfer and Bi-elliptic transfer and comparing between Hohmann transfer and Bi-elliptic transfer for long distance. It has been concluded that Hohmann transfer is more appropriate when the ratio of radius of final orbit to initial orbit () is less than 11.94. The purpose of this work is to minimize total full requirements, as well known that no refueling station in space, then using the computed ∆ʋ for determining the mass propellant consumed , at different specific impulse of the propellants, help us to carefully plane a mission to minimize the propellant mass carried on the rocket.

No abstract

The solar radiation pressure is one of the non-conservative force perturbation effects on earth satellite. A method presented to calculate the effect of radiation pressure on orbital elements for different altitude of satellites, the necessary calculation down by using modern Mat lab computer language. Typical results presented for radiation effect at deferent values of area mass ratio(A/m), the computed results indicated that the orbital elements much affected by solar radiation at high value of area mass ratio, also the effect on high orbit satellite is much more than that on low orbit satellite. The study also include the solar radiation pressure at two different value of ascending node ( Ω = 00 and Ω = 1800 ).

The main objective of this paper is to calculate the perturbations of tide effect on LEO's satellites . In order to achieve this goal, the changes in the orbital elements which include the semi major axis (a) eccentricity (e) inclination , right ascension of ascending nodes ( ), and fifth element argument of perigee ( ) must be employed. In the absence of perturbations, these element remain constant. The results show that the effect of tidal perturbation on the orbital elements depends on the inclination of the satellite orbit. The variation in the ratio  decreases with increasing the inclination of satellite, while it increases with increasing the time.

Researching and modeling perturbations is essential in astrodynamics because it gives information on the deviations from the satellite's normal, idealized, or unperturbed motion. Examined the impact of non-conservative atmospheric drag and orbital elements of low-earth-orbit satellites under low solar activity. The study is consisting of parts, the first looks at the effects of atmospheric drag on LEO satellites different area to mass ratios, and the second looks at different inclination values. Modeling the impacts of perturbation is included in each section, and the final portion determines the effects of atmospheric drag at various node values. The simulation was run using the Celestial Mechanics software system's SATORB module (Beutler, 2005), which solves the perturbation equations via numerical integration. The findings were examined using Matlab 2012. Conclusion that the impacts are stronger for retrograde orbits, which is due to the fact that the satellite moves in the opposite direction. The atmospheric drag effects for all orbital elements were increased by increasing the area to mass ratio. When the node value rises, the size parameter changes slightly, but the other orbital elements change. At varying inclinations, it is found that the changes in orbital elements due to atmospheric drug.