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
The influence on the regression rate of the inlet parameters of a solid fuel (PMMA) under supersonic cross flow has been investigated theoretically and numerically. Based on two-dimensional compressible N-S equations and species transport equations, the regression rate values were obtained by adding to the source term in the quality、 momentum and energy of the fuel. With different inlet parameters , i.e., stagnation temperature and pressure , the combustion conditions were analyzed at the flame-holding zone and the constant diameter cylindrical section. Comparison between numerical results with test data showed fair agreement, indicating the correctness of the calculation model. The results show that increasing the amount of oxygen or the mixture of the fuel and oxidant can improve the combustion efficiency; the inlet mass flow of air、stagnation temperature and depth of cavity has a great impact on the regression rate, then the impact of the length of cavity was slightly smaller, the impact of the stagnation pressure was minimum.
NomenclatureP =pressure T =temperature u, v =axial, radial velocity h =enthalpy value =density M =Mach number r =regression rate D =diameter L =length w q =heat flux subscripts fh =flameholding zone cyl =cylindrical section
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