Development of a Model Scramjet Combustor

Abstract

A model scramjet combustor has been developed to reproduce the main flow features that are present in a scramjet engine on a typical hypersonic vehicle; these flow features include shock/boundary layer interactions, shock/jet interactions, and supersonic mixing and combustion. The model combustor has been designed with optical access on three sides via fused-silica windows, and has a single fuel injector on the streamwise centerline capable of injecting fuel perpendicular to the bottom wall. The top wall of the inlet of the model combustor is ramped, with a turning angle of 10 ◦ , which generates a shock train in the model. We subject the model combustor to flows representative of the aerothermodynamic conditions expected in a scramjet combustor (Mach number, M ∼ 2.8, static pressure, P ∼ 40 kPa, temperature, T ∼ 1250 K) using the Stanford 6-inch Expansion Tube Facility. Fuel (hydrogen) is injected transversely into the freestream (air) with a momentum flux ratio J of 2.5, yielding an overall equivalence ratio φ of 0.25. Static pressure measurements on the top wall of the model have been made, and high-speed schlieren images have been acquired to characterize the flow. Additionally, OH* chemilluminescence images have been acquired to identify the burning region. We present the development, design, and features of the combustor, followed by an overview of measurements made thus far, and we conclude with a summary of our future plans.