Abstract: A rocket engine comprises first and second combustion chambers with respective combustion chamber liners bounding respective annular passages, wherein the first combustion chamber discharges into the second, and the respective annular passages are in fluid communication with one another. A portion of the effluent from the first combustion chamber flows from the first combustion chamber to the second combustion chamber through the respective annular passages via orifices in the respective combustion chamber liners, so as to provide for effusion cooling of a surface of the second combustion chamber. The first combustion chamber preferably operates fuel rich, reducing the temperature of the effusion cooling gases, which may be further cooled by a portion of unburned fuel. A flow restriction such as a turbine between the first and second combustion chambers provides a pressure differential therebetween that induces flow of effusion cooling gases.

Abstract: A rocket engine comprises first and second rotary injectors for injecting respective fuel and oxidizer propellant components into a first combustion chamber, and the effluent therefrom drives a turbine that rotates the rotary injectors. The mixture within the first combustion chamber is preferably fuel-rich so as to reduce the associated combustion temperature, and the fuel-rich effluent mixes in a second combustion chamber with additional oxidizer injected by a third rotary injector so as to generate a high temperature effluent suitable for propulsion. The rotary injectors are adapted so as to isolate the low pressure propellant supply from the relatively high pressures in the respective combustion chambers.