Investigation on the Effect of Thrust Chamber Geometric Parameters on the Performance of NTO/MMH Thrusters

Abstract

Hypergolic bipropellant radiation cooled thrusters utilizing Mono-methyl Hydrazine and Nitrogen Tetroxide are commonly used in spacecraft missions for attitude and orbit control. The performance index of a rocket engine is the specific impulse which is a function of combustion efficiency, known as C* efficiency, and the nozzle efficiency. An experimental investigation is carried out to evaluate the effect of combustion chamber design parameters on performance(C*efficiency) for varied injector spray and atomisation characteristics as well as injection pressures. Analytical model with empirical correlation available in the literature is used as a tool for understanding the physical process in the combustion chamber and predicting C* efficiency which was validated with experiments. Design variables considered are characteristic length and the contraction ratio. Cold flow evaluation of the injector using simulant water was done to evaluate the droplet size and injection velocity, which is normalised to the propellant flow conditions. Hot test for 10s using the stainless steel chamber was done at sea level with instrumentation for chamber pressure, mass flow rate of propellants and throat temperature. Results show that for a given injection and operating conditions, there exists a range of L* and contraction ratio where C*efficiency will be optimum and less sensitive. Trends in throat temperature measured are also evaluated. This paper presents the details of the investigation.