Dual-mode combustion of hydrogen in a Mach 5, continuous-flow facility

Results of an experimental and numerical study of a dual-mode scramjet comb ustor are reported. The experiment consisted of a direct-connect test of a Mach 2 hydrogen-air combustor with a single unswept-ramp fuel injector. The flow stagnation enthalpy simulated a flight Mach number of 5. Measurements were obtained using conventional wall instrumentation and a particle-imagi ng laser diagnostic technique. The particle imaging was enabled through the development of a new apparatus for seeding fine silicon dioxide particles into the combustor fuel stream. Numerical simulations of the combustor were performed using the GASP code. The modeling, and much of the experimental work, focused on the supersonic combustion mode. Reasonable agreement was o bserved between experimental and numerical wall pressure distributions. How ever, the numerical model was unable to predict accurately the effects of c ombustion on the fuel plume size, penetration, shape, and axial growth.