Computational Analysis of Straight Nozzle: Technical Note

Abstract

Most of the modern aircraft and military aircraft are powered by the modern gas turbine engine. They have nozzles to produce the required speed. Depending upon the required exit Mach number, a nozzle can be designed to be used for subsonic and supersonic flows. For the sonic flows, the convergent nozzle is used and for supersonic flows a convergent–divergent (CD) nozzle is used. In a CD nozzle, a straight nozzle flow is accelerated from low subsonic to sonic velocity at the throat and further expanded to supersonic velocities at the exit. This paper focuses on designing a straight nozzle to attain super-sonic flow and optimizing it to achieve maximum thrust without flow separation due to shock waves. This research also confirms that at which angle of deflection on the divergent portion produces more speed. The flow conditions were selected in view of the pressure, temperature and gases that are accessible at the exit of the combustion chamber. At the exit of the nozzle, the shock induced flow separation due to, over, under and optimum expansion conditions were studied.

Stage reduced counter-rotating axial compressor for jet engine

“…Nozzle inlet temperature (T 04 ) is calculated from the mechanical efficiency formula by substituting the values of turbine inlet temperature (T 03 ), the mass flow rate of fuel (M f ) and air (Ma), mechanical efficiency (h m ), and compressor outlet temperature (T 02 ) [19,20].…”