The cryogenic rocket engine using propellants of liquid oxygen (LOX) and also the liquid hydrogen (LH2) category as produces the highest specific impulse has been designed, constructed which comprises of a pressure feed system, injector head assembly thrust chamber, and conical nozzle. This proposed paper describes the design consideration of the cryogenic rocket engine is 1 tonne thrust level low temperature propellants LOX/LH2 mixture ratio (O/F) 5.2 with regenerative cooled rocket engine. The chamber pressure about 35kg/cm2 . The design philosophy for cryogenic rocket engine parameters and dimensions were calculated along with standardized equation like continuity and isentropic flow equations. The various components of cryogenic rocket performance are calculated using Rocket Performance Analysis (RPA) software like altitude performance, thermodynamic properties, fractions of the combustion products, estimated delivered performance etc. The pressure and temperature distribution inside the thrust chamber and nozzle 2D geometry by GAMBIT and a CFDFLUENT toolis used for investigation. The device fabrication has been done and the results were interpreted with the theoretical model.