ABSTRACT
A large rocket engine for the booster stage of a launch vehicle is being designed. This engine is to deliver a sea-level thrust of 1,500,000 1b f. The propellants are to be liquid oxygen (LOX) and RP-1. (The designation RP-1 refers to a grade of kerosene that is quite highly refined, particularly with regard to its sulfur content.) The engine chamber pressure is chosen as 1000 psia. The required propellant flow rates are 1700 lb m / s of RP-1 and 4070 lb m/s of LOX. These propellant are to be displaced by a turbopump. There will be a hot-gas generator employed for turbine drive, and the above-quoted flow rates include the anticipated gas generator propellant flows. We have been requested to provide a design for the turbine component of the engine's turbopump system.
