A new kind of power cycle of the propellant feed system for liquid propellant rocket engine is presented.The full flow staged combustion cycle and its superior performance compared to the traditional cycle are discussed.For further study of this advanced cycle system

experimental equipment for small scale rocket engine with full flow staged combustion cycle was designed and established.One-dimensional propellant feeding pipe system calculation has been carried out to confirm the pressure and volume needed for the propellant store tank.Through the thermodynamic calculation of two pilot burners

the burning temperature and the thermodynamic properties of the burning gas were obtained for further design of the pilot burner.A large number of specific impulse for the hydrogen-oxygen propellant was investigated with the chamber pressure and mixture ratio varying in a wide range.According to the performance analysis done

the chamber pressure and mixture ratio were determined to be 4.0 MPa and 0.75 respectively with certain consideration of the practical experimental conditions.Finally

the thrust generated was predicted to be 4 018.77 N.