The Effect of Fuel Injection Location on Supersonic Hydrogen Combustion in a Cavity-Based Model Scramjet Combustor

Eunju Jeong,Sean O’Byrne,A F P Houwing,In-Seuck Jeung

doi:10.3390/en13010193

Abstract

Supersonic combustion experiments were performed using three different hydrogen fuel-injection configurations in a cavity-based model scramjet combustor with various global fuel–air equivalence ratios. The configurations tested were angled injection at 15° to the flow direction upstream of the cavity, parallel injection from the front step, and upstream injection from the rear ramp. Planar laser-induced fluorescence of the hydroxyl radical and time-resolved pressure measurements were used to investigate the flow characteristics. Angled injection generated a weak bow shock in front of the injector and recirculation zone to maintain the combustion as the equivalence ratio increased. Parallel and upstream injections both showed similar flame structure over the cavity at low equivalence ratio. Upstream injection enhanced the fuel diffusion and enabled ignition with a shorter delay length than with parallel injection. The presence of a flame near the cavity was determined while varying the fuel injection location, the equivalence ratio, and total enthalpy of the air flow. The flame characteristics agreed with the correlation plot for the stable flame limit of non-premixed conditions. The pressure increase in the cavity for reacting flow compared to non-reacting flow was almost identical for all three configurations. More than 300 mm downstream of the duct entrance, averaged pressure ratios at low global equivalence ratio were similar for all three injection configurations.

Highlights

Scramjet engines are certainly an attractive propulsion system for next-generation, high-speed aircraft
As this paper is an extension of these previous studies, the focus here is on experimentally studying the influence of three different fuel injection locations: angled injection upstream of the cavity, parallel injection from the front face of the cavity, and upstream injection from the rear face of the cavity, at Mach 4 combustor inlet flow conditions
This paper investigated how combustion phenomena changed when the parameters of fuel This paper investigated how combustion phenomena changed when the parameters of fuel injection location, equivalence ratio, and total enthalpy of test gas were varied in a model supersonic injection location, equivalence ratio, and total enthalpy of test gas were varied in a model supersonic combustor with a cavity, using qualitative OH planar laser-induced fluorescence visualization and combustor with a cavity, using qualitative OH planar laser-induced fluorescence visualization and time-resolved floor static pressure measurements

Summary

Introduction

Scramjet engines are certainly an attractive propulsion system for next-generation, high-speed aircraft. As the fuel jet interacts with the supersonic crossflow, a bow shock is generated and fuel–air mixing, ignition, and combustion occur in the separation region. The wake region downstream of the barrel shock enhances fuel–air mixing This injection method generates stagnation pressure losses due to the strong bow shock generated by the transverse jet [1,2,3,4,5]. As this paper is an extension of these previous studies, the focus here is on experimentally studying the influence of three different fuel injection locations: angled injection upstream of the cavity, parallel injection from the front face of the cavity, and upstream injection from the rear face of the cavity, at Mach 4 combustor inlet flow conditions

Cavity-Based Model Scramjet Combustor

Free-Piston Shock Tunnel and Flow Conditions

OH PLIF Imaging

Pressure Measurements

Conclusions