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Numerical simulations were performed to analyze the performance of solid fuel scramjet (SFSCRJ) operating at Mach 4 to 6, with the primary goal being to investigate the feasibility of low takeover speed for SFSCRJ. The numerical model was created based on previous experiments with a constant area isolator added before combustor. Airflow is changed to subsonic in isolator and return to supersonic in the diverging section at the Mach 4 condition. The flow was not thermally choked at higher Mach condition, with a series of compression and expansion waves generating in the flameholding zone. The effect of flight Mach number is discussed from a viewpoint of total pressure recovery and combustion efficiency. The results show the feasibility of SFSCRJ operating at low flight Mach. To investigate the effect of inflow condition on combustor performance, different contraction ratios in the inlet are assumed at flight Mach 5 condition, resulting in different pressures at the entrance of isolator for numerical simulations. It is found that the performance of combustor is enhanced with the increase of inflow static pressure. There exists a tradeoff between the improved performance of combustor and the increased total pressure loss in the inlet. Nomenclature P = pressure T = temperature Y = mass fraction u = velocity H = isolator radius  = density M = Mach number f = friction force m  = mass flow rate F i = impulse function at the entrance of the isolator F e = impulse function at the exit of the isolator  = combustion efficiency

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The specific thrust level and variation rule of solid fuel scramjet combustor were researched both by numerical and experimental methods. A methane-burning vitiated air heater was designed and made to simulate flight environment of Mach 6 at 25km altitude. Based on the two dimensional flow field structures, a new quasi-one dimensional numerical method was established. The typical unsteady matter of solid fuel scramjet combustion and flow was transformed into a steady calculation of every moment in this method. And it can be used to simulate the parametric changing process of combustor quickly. Self-ignition and fuel regression rate characteristics in current experiment were consistent with results from previous works. And the agreement between the numerical and experimental results was generally good too. The specific thrust can reach the level of (600 ~ 800) N/kg· s in the designed flight environment. It was found that the specific thrust decrease with the total pressure loss increase during the working process. NomenclatureA = area B = pre-exponential factor of solid fuel pyrolysis c p = specific heat at constant pressure c ps = specific heat of solid fuel d = diameter dA = unit area increments D e = hydraulic diameter f dF = unit friction dm = unit mass increments f dQ = unit heat increments f f = friction coefficient F = thrust h = convective heat transfer coefficient H = flight altitude h g = effective vaporization heat m = mass flow rate Ma = Mach number f Nu = Nusselt number Pr = Prantdtl number p = pressure q = effective combustion heat of solid fuel Q = The low calorific value of solid fuel 1 PhD student, School of Aerospace Engineering, 2 r = fuel regression rate L = perimeter t = combustor working time T f = fluid temperature near the wall T so = initial temperature of solid fuel v = velocity Greek  = dynamic coefficient of viscosity  = heat conductivity coefficient of fluid  = density  = combustion efficiency  = gradient Subscripts a = ground atmospheric environment ground = ground test condition e = at the exit of combustor f = fluid flight = flight environment parameters i = time coordinate point ideal = ideal condition j = axial coordinate point s = solid phase sp = specific w = wall 0 = total parameter