Experimental Investigation of Supersonic Ramjet Combustion (Scramjet)
W. Waidmann, U. Brummund, Joachim Nuding
Abstract
Due to the importance of supersonic combustion in hypersonic propulsion, the flowfield in a supersonic ramjet combustion chamber has been experimentally investigated. Hydrogen was injected parallelly through a symmetric wedge into a heated, Mach 2.0 air stream. The wedge was centered as a strut in the two - dimensional combustion chamber and served as a flameholder. After ignition, a stable flame was established just downstream of the wedge, even at air temperatures where autoignition does not occur. Corresponding to the realized maximum air temperature of about 1500 K, a free flight Mach number up to M = 5.5 was simulated. Several measurement methods, schlieren pictures, Rayleigh - scattering images, OH - emission, laser - induced - predissociation - fluorescence (LIPF) and laser - doppler - velocimetry (LDV) have been used for characterization of the flow field and the combustion. The gas dynamic and combustion process including the interaction of shock waves with the mixing layer and the combustion zone will be discussed. The shock wave causes an intensification of combustion and an improved mixing. The combustion zone can be divided into an induction, transitional and highly turbulent region. Large - scale structures originate in the transitional region with an intensive momentum exchange. With combustion the velocity and Mach number in the combustion zone decreases to low subsonic values. The transverse growth rate of the flame strand is limited. The convective Mach number is about M = 1.6.