As a preliminary step for the research and development of a bipropellant thruster using methane as fuel, the following study was conducted as follows; an analysis of the theoretical performance characteristics for methane-fuel bipropellant small rocket engine, an injector performance evaluation, and combustion experiments of the gaseous methane and oxygen. Attention has been drawn to the need for reusable rockets in order to reduce the dramatic cost of space propulsion engine development, and a demand for eco-friendly propellants research has also increased. Especially, liquid methane has been superior in terms of performance/economy compared to kerosene and liquid hydrogen which are currently used as fuels for space launch vehicles. Moreover, Potentially the biggest advantage of using methane is that this fuel exist abundantly or can be synthesized on many planets that future spacecrafts will visit. Thus, it is considered relevant as a high-performance/low-cost propellant. And design parameters required for methane-oxygen bipropellant small-rocket-engine were derived through the theoretical performance analysis. The theoretical performance of the rocket engine was analyzed by using CEA and optimal propellant mixture ratio, characteristic length, and expansion ratio of a nozzle were calculated by assuming chemical equilibrium. the performance evaluation of the injector was carried out for that is one of the critical components of bipropellant-rocket-engine. Spray characteristics are investigated in detail according to the recess ratio and injection pressure on the swirl-coaxial injector using simulant propellants such as gaseous nitrogen and water. The spray angle and breakup length were measured based on the shape of the spray from the swirl-coaxial injector, and the spray characteristics were analyzed based on whether mixing occurred inside or outside of the injector. The recess ratio has an impact on the spray characteristics by changing the mixing location, and thus should be considered an important design factor in relation to injector performance. And combustion experiments were performed under various propellant injection conditions to analyze the combustion characteristics of methane-oxygen diffusion flames in a model combustion chamber, thereby measuring the combustion stability limit and observing the flame shape variation. The combustion stabilization study was analyzed through image post-processing from direct image of flame shape, and the effect of flame structure and recess of the injector on combustion was investigated by visualization of spontaneous OH radical chemiluminescence. Also, a wide and short flame was created by the swirl effect of the propellant, indicating a suitability of the swirl-type injector when it is employed in the thruster with a limited length of combustion chamber.