Recent space developments are implementing several simpler and less expensive rocket technologies. Environmental concerns and following governmental restrictions necessitate to replace current (hydrazine-based) toxic propellants with green ones, with a minimum loss of performance. Hydrogen peroxide is a promising candidate for the future of green propellants due to its flexibility and its benign nature allows the advancement of simple, cost-effective, and environmentally friendly propulsion with sufficient performance to replace hydrazine or other highperforming toxic propellants. Consequently, this thesis is devoted for the study of hydrogen peroxide-based propellants for future space propulsion applications. The main objective of this work is to study the combustion properties of green propellants. Foremost we discussed the hydrogen peroxide use, properties, and management of in-space propulsion, and later, various combinations and compositions with hydrogen peroxide as a bipropellant have been studied using NASA CEA code. The main purpose is to find combustion temperature and specific impulse values at different O/F ratios of 2,4,6,8,10 and various pressure chamber values of 20, 25, and 30 bar. For this purpose, two cases have been considered to study the bi propellant of liquid methane and mass fraction variation is obtained at different O/F ratio and at chamber, throat and exit. Analysis has done considering all the compositions and comparison of combustion products in the case of bi propellant in order to achieve the best efficiency at proper O/F ratio and fixed chamber pressure. It is observed that concentration of hydrogen peroxide has the significant effect on combustion performances and the chemical composition effects due to weight concentration. It is concluded that hydrogen peroxide is useful for the future development of the research activity.

It was realized earlier that chemical propulsion systems utilize fuel very inefficiently, which greatly limits their lifespan. Electric propulsion is into existence to overcome this limitation of chemical propulsion. The magnetoplasmadynamic (MPD) thruster is presently the most powerful form of electromagnetic propulsion. It is the thruster’s ability to efficiently convert MW of electric power into thrust which gives this technology a potential to perform several orbital as well as deep space missions. MPD thruster offers distinct advantages over conventional types of propulsion for several mission applications with its high specific impulse and exhaust velocities. However, MPD thruster has limitations which limits its operational efficiency and lifetime. In this paper, the thruster limitations are reviewed with respect to three operational limits i.e., the onset phenomenon, cathode lifetime, and thruster overfed limits. The dependence and effects of the operational limits on each other is compared using different empirical models to derive a scaling factor that has been found for each geometrical arrangement; a limiting value exists beyond which the operation becomes highly unsteady and electrode erosion occurs. Along with reviewing and proposing methods to overcome power limitations for MPD thrusters, the relation between exit velocity and ratio of electrode’s radius is also verified using Maecker’s formula.