def solve(config, t, R, mdot_ox, mdot_fuel, Pc, OF, Cstar, Pe, thrust, ISP):
rho_fuel = config.getfloat('Params', 'rho_fuel')
L = config.getfloat('Params', 'fuel_length') / M_TO_IN
a = config.getfloat('Params', 'a')
n = config.getfloat('Params', 'n')
Rt = config.getfloat('Params', 'throat_radius')
AR = config.getfloat('Params', 'area_ratio')
P0 = config.getfloat('Params', 'p0')
throat_area = np.pi * Rt**2
N = t.shape[0]
for i in range(N-1):
dt = t[i+1] - t[i]
R[i+1] = advance_R(a, n, R[i], mdot_ox[i], dt)
mdot_fuel[i+1] = utils.get_mdot_fuel(R[i+1], L, rho_fuel,
mdot_ox[i+1], a, n)
mdot = mdot_fuel[i+1] + mdot_ox[i+1]
Pc[i+1] = Cstar[i] / throat_area * mdot
OF[i+1] = mdot_ox[i+1] / mdot_fuel[i+1]
gamma, _, Cstar[i+1] = utils.get_CEA_results(CEA_object, Pc[i+1],
P0, OF[i+1], AR)
Pe[i+1] = utils.get_exit_pressure(Pc[i+1], gamma, AR, guess=Pe[i])
CF = utils.get_thrust_coeff(Pc[i+1], Pe[i+1], P0, gamma, AR)
thrust[i+1] = Pc[i+1]*throat_area*CF
ISP[i+1] = thrust[i+1] / G0 / (mdot_ox[i+1] + mdot_fuel[i+1])
def initialize(config, R_0, mdot_ox_0, Cstar_0):
n_timesteps = config.getint('Params', 'n_timesteps')
burn_time = config.getfloat('Params', 'burn_time')
P0 = config.getfloat('Params', 'p0')
rho_fuel = config.getfloat('Params', 'rho_fuel')
L = config.getfloat('Params', 'fuel_length') / M_TO_IN
a = config.getfloat('Params', 'a')
n = config.getfloat('Params', 'n')
Rt = config.getfloat('Params', 'throat_radius')
AR = config.getfloat('Params', 'area_ratio')
throat_area = np.pi * Rt**2
t = np.linspace(0, burn_time, num=n_timesteps+1)
R = np.zeros(t.shape)
R[0] = R_0
mdot_ox = np.zeros(t.shape)
mdot_ox[0] = mdot_ox_0
mdot_fuel = np.zeros(t.shape)
mdot_fuel[0] = utils.get_mdot_fuel(R[0], L, rho_fuel, mdot_ox[0],
a, n)
Pc = np.zeros(t.shape)
Pc[0] = Cstar_0 / throat_area * (mdot_ox[0]+mdot_fuel[0])
OF = np.zeros(t.shape)
OF[0] = mdot_ox[0] / mdot_fuel[0]
Cstar = np.zeros(t.shape)
thrust = np.zeros(t.shape)
Pe = np.zeros(t.shape)
ISP = np.zeros(t.shape)
gamma, _, Cstar[0] = utils.get_CEA_results(CEA_object, Pc[0], P0,
OF[0], AR)
Pe[0] = utils.get_exit_pressure(Pc[0], gamma, AR)
CF = utils.get_thrust_coeff(Pc[0], Pe[0], P0, gamma, AR)
thrust[0] = Pc[0] * throat_area * CF
ISP[0] = thrust[0] / G0 / (mdot_ox[0] + mdot_fuel[0])
return t, R, mdot_ox, mdot_fuel, Pc, OF, Cstar, Pe, thrust, ISP