# e) answers
print('stagnation pressure [bar]: ', p_0i)
Qm = u_e * rho_e * b * 2 * h_duct #kg/s
print('massflow [kg/s]', Qm)
"""
2) compute M,p,p0 along the nozzle and the duct
"""
po_star = p_0_star_e
#a) nozzle
x = np.linspace(0, 120, 1000)
ones = np.ones(len(x))
h = np.zeros(len(x))
for i in range(len(h)):
h[i] = nozzle_height(x[i]) / 1000
air_ratios = h_throat / h
Mx = useful.inv_air_ratio(air_ratios, 0.5 * np.ones(len(x)))
px = p_0i / ((1 + 0.2 * Mx**2)**(1.4 / 0.4))
np.savetxt("results\homework_a_nozzle.txt", np.stack((x, Mx, px, p_0i * ones)))
# fig, ax = plt.subplots()
# ax.set_title('a nozzle')
# ax.plot(x,Mx,'-b',label='M')
# ax.plot(x,px,'-r',label='p[bar]')
# ax.plot(x,p_0i*ones,'-c',label='p0[bar]')
# ax.legend()
# ax.grid()
#b) duct
x2 = np.linspace(0, l_duct, 1000)
l_nozzle = 0.12 # [m] l_duct = 1.5 # [m] xe = l_duct # [m] de = 2 * h_duct dt = 0.006 # 6 [mm] b = 0.05 # 50 [mm] Dh_duct = 4 * h_duct * b / (b + 2 * h_duct) # 4*A/P p_a = 1.01325 # 1.01325[bar] = 1[atm] T0 = 300 # [K] # Gas = air (assuming air as a perfect gas) gamma = 1.4 R = 287.1 # [J/kg/K] p_e = p_a x_sh = 0.07 # [m] h_sh = nozzle_height(x_sh * 1000) / 1000 #[m] """ 2) resolution de l exercice' """ # Throat conditions = sonic conditions T_star = 2 * T0 / (gamma + 1) c_star = np.sqrt(gamma * R * T_star) # Shock relations : # -1- Isentropic between reservoir and sh1 # -2- Isentropic between sh2 and end of the nozzle #1) determiner les conditions juste avant le shock M_sh1 = np.array([1.5]) air_ratio1 = h_throat / h_sh #ratio between the throat and the first shock
import numpy as np from nozzle import nozzle_height import useful from matplotlib import pyplot as plt #données de base # hello world h_throat = 0.003 #3mm h_duct = 0.0075 #7.5 mm length_nozzle = 0.12 length_duct = 1.5#58.29722 d = 2*h_duct b = 0.05 #50mm Dh_duct = 4*h_duct*b/(b+2*h_duct) #4*A/perimeter x_shock= 0.07 h_shock = nozzle_height(x_shock*1000)/1000 #[m] p_a = 1.01325 #btm gamma = 1.4 R=287.1 #J/kg/K T0 = 300 #K #b.1) Determine the sonic conditions A_t = A* T_star = T0/(gamma+1)*2 c_star = np.sqrt(gamma*R*T_star) #b.2) determiner les conditions juste avant le shock air_ratio1 = h_throat/h_shock #ratio between the throat and the first shock M1 = useful.inv_air_ratio_sup(air_ratio1,np.array([1.5])) T1 = T0/(1+0.2*M1**2) u1= M1*np.sqrt(1.4*R*T1)
l_nozzle = 0.12 # [m]
l_duct = 1.5 # [m]
xe = l_duct # [m]
de = 2 * h_duct
dt = 0.006 # 6 [mm]
b = 0.05 # 50 [mm]
Dh_duct = 4 * h_duct * b / (b + 2 * h_duct) # 4*A/P
p_a = 101325.0 # [Pa] = 1 [atm]
T0 = 300 # [K]
# Gas = air (assuming air as a perfect gas)
gamma = 1.4
R = 287.1 # [J/kg/K]
p_e = p_a
x_sh = 0.12 # [m]
h_sh = nozzle_height(x_sh * 1000) / 1000 #[m]
# Throat conditions
T_t = 2 * T0 / (gamma + 1)
T_star = T_t
c_t = np.sqrt(gamma * R * T_t)
u_t = c_t
# Shock relations :
# -1- Isentropic between reservoir and sh1
# -2- Isentropic between sh2 and end of the nozzle
M_sh1 = 1.5
M_sh1, error_M_sh1, iter_M_sh1 = fun.ratio_A_star_over_A_super(
h_throat, h_sh, M_sh1)
print('M_sh1 = ', M_sh1)
T_sh1 = T0 / (1 + 0.5 * (gamma - 1) * M_sh1**2)
from nozzle import nozzle_height import useful from matplotlib import pyplot as plt """ données de base """ h_throat = 0.003 #3mm h_duct = 0.0075 #7.5 mm length_nozzle = 0.12 l_duct = 1.5 #58.29722 de = 2 * h_duct b = 0.05 #50mm Dh_duct = 4 * h_duct * b / (b + 2 * h_duct) #4*A/perimeter x_shock = 0.12 h_shock = nozzle_height(x_shock * 1000) / 1000 #[m] pa = 1.01325 # 1 atm = 1.01325 bar gamma = 1.4 R = 287.1 #J/kg/K T0 = 300 #K #1) Determine conditions at the throat = sonic conditions A_t = A* T_star = T0 / (gamma + 1) * 2 c_star = np.sqrt(gamma * R * T_star) # Shock relations : # -1- Isentropic between reservoir and sh1 # -2- Isentropic between sh2 and end of the nozzle #2) determiner les conditions juste avant le shock air_ratio = h_throat / h_shock #ratio between the throat and the first shock