Two-stream supersonic turbulent mixing layers with a free-stream Mach number of 2.3 on the high-velocity side are experimentally investigated. A large-scale vortical structure, which has been believed to dominate the development of incompressible mixing layers, is also observed in the present supersonic layers. The spreading rate of the layer is correlated with a velocity ratio of the free streams and a Mach number based on the velocity difference across the layer.
H 2 -fueled scramjet engines were tested under Mach 4 (M4) to Mach 8 (M8) ight conditions, and the local equivalence ratio and combustion ef ciency were measured by gas sampling at the engine exit. Correlation between the local values of equivalence ratio and combustion ef ciency showed that the M4 combustion was principally reaction controlled and the reaction and the mixing-controlled combustion coexisted in the M6 condition. The M8 combustion in the engines was rate controlled by the mixing of H 2 . Comparison of engine performance in the air, supplied by a combustion heater and a storage heater, indicated that the performance strongly depended on the air-heating methods. The dependence of engine performance on air-heating methods could be explained by the nding that the M6 engine combustion was partially reaction controlled. The wall-heating rate and pressure distribution in the M6 tests also supported the shift from the partially reaction-controlled to the mixing-controlled combustion as the fuel rate was increased. The mixing-controlled combustion suggested weak facility dependence in the M8 condition. Because reaction in the M8 condition is not suf ciently fast, the main combustion region might be blown downstream in the engine. NomenclatureH = height of the engine (250 mm) k 1 = reaction rate of the reaction H C O 2 ! OH C H M = Mach number n = pressure exponent in Eq. (3) (order of overall reactions) P b = typical burning (static) pressure weight-averaged with n D 1:6 P i = typical ignition (static) pressure weight-averaged with n D 1 P w = wall pressure S = heated by a storage air heater T r = recovery temperature in engines t b = burning time of H 2 in engines (assumed to be t 95 ) t ow = residence time in the burning length, x e t ig = ignition time de ned as 10 £ t 1 t s = typical gas sampling time (1.5 s) t 1 = characteristic time de ned as (k 1 [O 2 ]) ¡1 t 95 = reaction time to raise the temperature from 5 to 95% of the equilibrium temperature rise U = typical ow velocity in engines V = heated by a vitiation air heater x b = characteristic length required to burn fuel x e = burning length in engines (1.3 m) x ig = characteristic length required to ignite fuel 8 = bulk fuel equivalence ratio supplied to engines Á = local fuel equivalence ratio (time-averaged, if not speci ed) c = local combustion ef ciency (time-averaged, if not speci ed) -= time-or space-averaged values
Thrust performances of scramjet engines were compared with theoretical values to quantify the progress in engine performances from Mach (termed as "M") 4 to M8 flight conditions. An engine with a ramp produced a net thrust of 215 N under the M8 tests and a comparison of a theoretical thrust yielded a thrust achievement factor of 51%. By excluding boundary layer, an engine with a thick strut delivered a net thrust of 560 N and showed a thrust factor of 92% and a net thrust factor of 45%. The thrusts were limited by flow separation caused by engine combustion (termed as "engine unstart"). The starting characteristics was improved by boundary layer controls in M6 and M4 conditions. An engine with a thin strut doubled the thrust from 1620 N to 2460 N by the boundary layer bleeding in the M6 tests. The improved thrust factor was 60% at the stoichiometric H 2 condition. Under M4 tests, the net thrust was tripled by the bleed and a two-staged injection of H 2. As results, the thrust factor was raised from 53% to 70%, the net thrust factor was increased from 32% to 55%. Studies required for improving the net performance was addressed. Nomenclature A 1 Inlet area 0.2 m-wide 0.25 m-high) cf wall friction coefficient Cint Internal drag coefficient of engines Dint Internal drag of engines = Cint q 1 A 1) Df Internal friction drag on engines Df0 Minimum friction drag on the internal surface of a rectangular duct Dp Internal pressure drag of engines d1 Displacement thickness of boundary layer Isp Fuel specific impulse (km/s) M Mach number q 1 Freestream dynamic pressure at inlet (kPa) hp Total pressure recovery factor hc Air capture ratio in inlets DF Thrust increment by combustion DFnet Net thrust by combustion (= DF-Dint) hdrag Drag achievement factor (= Df0/Dint) hDF Thrust achievement factor (= DFexp /DF0) hnet Net thrust achievement factor e Geometrical contraction ratio of inlets F H 2 equivalence ratio subscripts exp experimental values 0 baseline values
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