Recent Developments in Rocket Nozzle Configurations

Abstract

T THRUST of a rocket engine is primarily dependent upon the momentum imparted to the products of combustion by discharging them through an exhaust nozzle. During their passage through the nozzle, the exhaust gases are continuously accelerated from low subsonic velocities to high supersonic velocities at the nozzle exit. For the purpose of understanding the flow process, the nozzle can be divided into three portions: the convergent subsonic section, the throat section, and the divergent supersonic section. This division is appropriate because of the different effects each part has in determining the thrust developed by the rocket, and because different methods of analysis have to be used in determining the flow field in these three regions. The mass flow rate through a rocket nozzle is determined by the area of the throat section and operating conditions of the combustion chamber. Design changes in the configuration of the convergent portion of the nozzle would influence the mass flow of the exhaust gases and also, to some extent, the combustion efficiency achieved in the chamber. The sonic velocity attained by the exhaust gases is fixed by the combustion chamber conditions and can be further increased by the expansion occurring in the divergent supersonic portion of the nozzle. This additional velocity, leading to thrust increment, is dependent only on the configuration of the diverging nozzle walls and the exit area. In recent years the design of the divergent supersonic portion of the rocket nozzle has received considerable attention. The scope of this review will be limited to this portion of the nozzle and improvement thereof. Analysis of rocket nozzle flows in any real case should include radiative heat loss, chemical reactions due to incomplete combustion, and chemical properties of the exhaust gases. However, gross performance comparisons can be made by assuming that the exhaust gases expand adiabatically and behave like an ideal gas with a constant ratio of specific heat capacities. The appropriate choice of nozzle configuration for a specific rocket engine may, to a great extent, depend on the fabrication methods of nozzle walls, cooling requirements, allowable limits on dimensions, influence of nozzle weight on overall rocket performance, etc. Detailed examination of all these features involves various special engineering fields aside from the knowledge of supersonic expansion process of the