• constellation
• deployment
• MATLAB
• mission analysis
• nanosatellite
• orbit
• spaceflight
• tool

The Master Thesis proposed has been done at D-Orbit, a company near Como (Italy). D-Orbit is famous for its in-orbit transportation service, InOrbit NOW (ION). ION consists of a carrier of CubeSats and it can deploy up to 48U. In this context, it is useful to have a semi-automatic MATLAB tool to do a preliminary mission analysis. In particular, the evaluation of the minimum propellant consumption, which guarantees the orbit transfers performed by ION to complete the deployment, is requested. In order to minimize the propellant consumption, the natural perturbations acting on a body orbiting Earth are exploited. The tool takes into consideration two direct transfers and the passage through three mission orbits, therefore in input there are the release orbit and three mission orbits of three different CubeSats. The direct transfer is performed from the release orbit to the first mission orbit. Next, the perturbations guarantee the reaching of the second mission orbit. The last orbit transfer is performed to reach the last mission orbit. In this way, at least three CubeSats can be released by the carrier. The inputs consist in the release orbit, provided by the launcher service provided, in the mission orbit of each CubeSat released by ION and in the time within which the deployment must be completed, these data are provided by the customer. The tool is conceived to be semi-automatic in such a way that the user has only to set the inputs and start it. Particular attention has been put on the effects of the oblateness of the Earth, also called J2 effect, and of the aerodynamic drag. Indeed, the first perturbation mainly induces a variation of the Right Ascension of the Ascending Node in time. The aerodynamic drag induces variations of the semi-major axis, it decreases in time, and the eccentricity, it goes to zero. These effects can be exploited to perform a transfer between two orbits which differ only in terms of RAAN, semi-major axis and eccentricity. The problem has been solved in two situations. The first situation implies the combination of the effects of the oblateness of the Earth and the aerodynamic drag while the second takes into account only the effect of the oblateness. The results and the requests of the company to have a tool capable to give the output in a few hours have led to choosing the simplest case, the one which considers only the J2 effect. Next, the minimization of the propellant consumption has been conducted using the MATLAB functions “fminunc” and “ga”, guaranteeing that the determined minimum is the global minimum and not local. The optimization has been done with respect to the orbital parameters of the first intermediate mission orbit, that is reached through the first direct transfer. The boundaries within which the variables can freely vary are imposed by the customers, as they usually give a margin of tolerance. To validate the tool is has been applied to a case, which solution was known. The results show that the method of optimization of the propellant consumption is valid and that further refinements are recommended to fine-tune the output value.