Vaporization of Propellants in Rocket Engines

Abstract

The theory and equations required to calculate the vaporization rates in a rocket engine, based on a model considering propellant vaporization as the rate controlling combustion process, are presented. Included in the model are changes in combustion gas velocity, droplet velocity, temperature and mass. The vaporization rates of heptane, ammonia, hydrazine, liquid fluorine and liquid oxygen sprays were computed for various log probability distributions and mass median drop sizes of the spray. The rates were also calculated for an engine operating with various injection velocities, final gas velocities, initial propellant temperatures and chamber pressures. The calculations of the per cent of propellant vaporized are correlated with an effective chamber length. The correlated results show tha t the percent of propellant vaporized can be increased either by decreasing the mass median drop size and the initial drop velocity or by increasing the chamber length, chamber pressure, final gas velocity and initial fuel temperature. A method of analysis tha t relates experimental rocket engine performance to the quantity of liquid propellant vaporized is also presented. With this method experimental engine performance results are compared with the calculations. The comparison suggests that incomplete vaporization is responsible for combustor inefficiency. Experimental results agree with the calculations for sprays having a geometric standard deviation of 2.3 and a mass median drop radius of 70 to 240 microns, depending on the type of injector.