Design of a Model Scramjet Combustor for Vortex-Enhanced Mixing and Combustion Studies

Abstract

The present work details the preliminary analysis and design of a model scramjet combustor intended to serve as a research platform for mixing and combustion studies in support of current non-reactive research. The shock tunnel located at the Aerodynamics Research Center of The University of Texas at Arlington, capable of producing test conditions simulating scramjet flight at Mach numbers ranging from ~5 to ~16 at altitudes of ~24 km to ~85 km, will serve as the test facility for the model combustor. Specifically, the modular design of the model scramjet combustor will allow for the study of supersonic combustion and flameholding characteristics of specific fuel injection strategies based on enhanced fuel-air mixing via pre-imposed selected modes of streamwise vortex interactions and consequent dynamics. In this paper, the fundamental requirements for the combustor design and the performance of the UTA hypersonic tunnel are discussed. A parametric sizing of the model combustor is utilized to define a suitable design that meets the specified requirements, and the inevitable constraints imposed by the limitation of ground testing facilities, discussed in the manuscript. Nomenclature= Area γ = Specific heat ratio h = Enthalpy M = Mach number P = Pressure ψ = Combustor to freestream flight static temperature ratio ρ = Density η = Adiabatic compression efficiency T = Temperature t = time θ = Inlet compression angle u = Velocity Subscripts 1 = Driven condition 4 = Driver condition 5 = Stagnation condition c = Combustor condition e = Nozzle exit condition i = ignition r = reaction * = Throat condition ∞ = Freestream condition