def __design_angelino_nozzle(self, design_alt, truncate_ratio, CEA, r_e):
""" Can be a class function, will edit going forwards
"""
(p_atm, T_atm, rho_atm) = gd.standard_atmosphere([design_alt])
PR = CEA.p_c / p_atm
M_e = gd.PR_expansion_mach(PR, CEA.gamma)
expansion_ratio = gd.expansion_ratio(1, M_e, CEA.gamma) #6.64 #8.1273
# print('Exp. ratio: ' + str(expansion_ratio))
# print('PR: ' + str(PR))
A_t = r_e**2 * np.pi / expansion_ratio # max expansion (r_b = 0, r_e**2 >= A_t*expansion_ratio/np.pi)
return plug_nozzle(expansion_ratio,
A_t,
r_e,
CEA.gamma,
CEA.T_c,
CEA.p_c,
CEA.a_c,
CEA.rho_c,
1500,
truncate_ratio=truncate_ratio)
def __init__(self,
design_alt,
r_e,
gamma,
T_c,
p_c,
a_c,
rho_c,
n,
truncate_ratio=1):
# input design parameters
(p_atm, T_atm, rho_atm) = gd.standard_atmosphere([design_alt])
PR = p_c / p_atm
M_e = gd.PR_expansion_mach(PR, gamma)
expansion_ratio = gd.expansion_ratio(1, M_e, gamma)
print('Expansion ratio: ' + str(expansion_ratio))
self.expansion_ratio = expansion_ratio
self.A_t = r_e**2 * np.pi / expansion_ratio
self.r_e = r_e
self.gamma = gamma
self.n = n
self.truncate_ratio = truncate_ratio
self.T_c = T_c
self.p_c = p_c
self.a_c = a_c
self.rho_c = rho_c
# calculated design parameters
self.A_e = self.A_t * self.expansion_ratio
self.r_b = np.sqrt(-self.A_e / np.pi + self.r_e**2)
self.M_e = optimize.fsolve(
lambda M: gd.expansion_ratio_zero(1, M, self.gamma, self.
expansion_ratio), 5)
# DESIGN OF NOZZLE, FUNCTION ORDER IS IMPORTANT
# NON-OPTIONAL FUNCTION RUNS
self.design_nozzle()
self.truncate_nozzle()
self.calc_flow_properties()
self.arc_length_coord()
# OPTIONAL FUNCTION CONSTANTS
self.converge_section = 0 # whether the converging section has been designed
def update_contour(self, x, y, M, x_centre_spike=0):
# Updates the spike contour with new (x,y) points with known values of M at each point
self.x = x
self.y = y
self.M = M
# optionally centre spike about x-axis
if (x_centre_spike):
self.centre_spike()
# update flow properties based on isentropic expansion
self.calc_flow_properties()
# update arc length coordinates
self.arc_length_coord()
# update exit mach number
self.M_e = M[-1]
# update expansion ratio
self.expansion_ratio = gd.expansion_ratio(1, self.M_e)
self.A_t = self.r_e**2 * np.pi / self.expansion_ratio
print(
"Warning, throat area update not complete, assumes perfect isentropic expansion from throat to exit"
)
# update area estimation
self.A = self.A_t * gd.expansion_ratio(1, self.M, self.gamma)
# update base radius
self.r_b = self.y[-1]
# update exit area
self.A_e = np.pi * (self.r_e**2 - self.r_b**2)
if (self.converge_section):
print(
"Warning, congerence section not updated. Run self.converge_section(args) again to define a new convergence section."
)
def design_nozzle(self):
# discrete contour design variables
self.M = np.linspace(1, self.M_e, self.n)
self.A = self.A_t * gd.expansion_ratio(1, self.M, self.gamma)
self.alpha = gd.prandtl_meyer(self.M_e, self.gamma) - gd.prandtl_meyer(
self.M, self.gamma) + gd.mach_angle(self.M)
self.l = (self.r_e - np.sqrt(
np.abs(self.r_e**2 -
(self.A * self.M * np.sin(self.alpha) / np.pi)))) / np.sin(
self.alpha)
self.x = self.l * np.cos(self.alpha)
self.y = self.l * np.sin(self.alpha)
self.centre_spike()
self.length = self.x.max()
def COST_FNC(design_alt, truncate_ratio, T_w, CEA, r_e, alpha, beta, n):
### DESIGNING NOZZLE
(p_atm, T_atm, rho_atm) = gd.standard_atmosphere([design_alt])
PR = CEA.p_c / p_atm
M_e = gd.PR_expansion_mach(PR, CEA.gamma)
expansion_ratio = gd.expansion_ratio(1, M_e, CEA.gamma) #6.64 #8.1273
# print('Exp. ratio: ' + str(expansion_ratio))
# print('PR: ' + str(PR))
A_t = r_e**2 * np.pi / expansion_ratio # max expansion (r_b = 0, r_e**2 >= A_t*expansion_ratio/np.pi)
spike = plug_nozzle(expansion_ratio,
A_t,
r_e,
CEA.gamma,
CEA.T_c,
CEA.p_c,
CEA.a_c,
CEA.rho_c,
n,
truncate_ratio=truncate_ratio)
### CALCULATING COST
## thurst estimation over altitude
alt_range = np.linspace(0, 12000, 30)
(p_atm_r, T_atm_r, rho_atm_r) = gd.standard_atmosphere(alt_range)
#print(CEA.p_c/p_atm_r)
thrust_range = np.zeros(alt_range.shape)
for i in range(alt_range.shape[0]):
if i == 10:
MOC_mesh = MOC.chr_mesh(spike,
gamma,
alt_range[i],
50,
downstream_factor=1.2,
plot_chr=0)
else:
MOC_mesh = MOC.chr_mesh(spike,
gamma,
alt_range[i],
50,
downstream_factor=1.2,
plot_chr=0)
thrust_range[i] = MOC_mesh.compute_thrust('nearest', 10)
work = np.trapz(thrust_range, alt_range)
plt.plot(alt_range, thrust_range, 'o')
plt.show()
## heat transfer required
total_heat_flux = heat_flux(CEA.Pr, CEA.cp, CEA.gamma, CEA.c, CEA.w,
CEA.T_c, T_w, spike)
# print('Work*alpha: ' + str(work*alpha))
# print('Heat flux*beta: ' + str(total_heat_flux*beta))
return -alpha * work + total_heat_flux * beta
design_alt = 9144
truncate_ratio = 1 # bounds on truncate < 0.1425
gamma = 1.237 #np.mean([1.2534,1.2852])
T_c = 2831.47 # combustion chamber temperature
p_c = 3102640.8 # combustion chamber pressure
rho_c = 3.3826 # combustion chamber density
a_c = np.sqrt(gamma * (1 - 1 / gamma) * 200.07 *
T_c) # combustion chamber sound speed
(p_atm, T_atm, rho_atm) = gd.standard_atmosphere([design_alt])
PR = p_c / p_atm
M_e = gd.PR_expansion_mach(PR, gamma)
expansion_ratio = gd.expansion_ratio(1, M_e, gamma) #6.64 #8.1273
# print('Exp. ratio: ' + str(expansion_ratio))
# print('PR: ' + str(PR))
A_t = r_e**2 * np.pi / expansion_ratio # max expansion (r_b = 0, r_e**2 >= A_t*expansion_ratio/np.pi)
plug1 = plug_nozzle(expansion_ratio,
A_t,
r_e,
gamma,
T_c,
p_c,
a_c,
rho_c,
1000,
truncate_ratio=1)
