Conceptual trust chamber design for upper stage liquid rocket engine

In this study, fuel, fuel mixture ratio (O/F), thrust chamber geometry and thrust chamber materials were conceptually determined for upper stage liquid rocket engine which can produce 2 kN thrust. N2O4 was selected as oxidizer and N2H4 was selected fuel. Considering performance criteria, 1.4 ideal fuel mixture ratio and 1 MPa combustion chamber pressure was specified by RPA (Rocket Propulsion Analysis) program analyzes. % 80 prabolic bell nozzle was designated as nozzle shape and DLR silca SF which is carbon-ceramic composite was selected as thrust chamber material. As a result of analyzes, specific impuls (Is) 338.7 s in the vacum, thrust coefficient (CF) 2.0018, characteristic velocity (c*) 1659.26 m/s and effective exhaust velocity 3321.55 m/s were calculated.