Experimental Investigation and Numerical Simulation of Secondary Combustor Flow in Boron-Based Propellant Ducted Rocket

Abstract

This paper establishes a two-phase reacting flow model in the secondary combustor of a ducted rocket. The threedimensional Favre-averaged compressible turbulent N-S equations are used as the governing equations of the reacting flow, the improved k $ " two-equation turbulence model is used to simulate the turbulent flow, and the eddy break up model is used to simulate the gas combustion. The particle-phase solution is obtained using a well-established boron particle ignition and combustion model. Boron particles are ejected from the exit of the gas generator into a secondary combustor and their trajectories are traced through the reacting flowfield using discrete phase models. The secondary combustion in a cylindrical combustor for the ducted rocket is investigated preliminarily with tests conducted on a connected-pipe ramjet test bed. During the test, boron-based fuel-rich HTPB propellant is used. The effect of factors, such as air/fuel ratio, velocity of air injection and dome height on the performance of the combustor and the engine is examined. The work described in this paper represents an attempt to direct the design of the secondary chamber in order to increase combustion efficiency. The experimental results show that the reacting flow model established in this paper is correct.