def __cost_end_func(self,
no_alt_range_int,
spike,
CEA,
downstream_factor=1.2,
chr_mesh_n=120,
no_core=1):
alt_range = np.linspace(0, 9144, no_alt_range_int)
# shuffle arrays so each core computes similar complexity on average
np.random.shuffle(alt_range)
(p_atm_r, T_atm_r, rho_atm_r) = gd.standard_atmosphere(alt_range)
thrust_range = self.__multicore_thrust_compute(spike,
alt_range,
CEA.gamma,
downstream_factor=1.2,
chr_mesh_n=chr_mesh_n,
no_core=no_core)
# unshuffle arrays
ordered_idx = np.argsort(alt_range)
alt_range = alt_range[ordered_idx]
thrust_range = thrust_range[ordered_idx]
work = np.trapz(thrust_range, alt_range)
# plt.plot(alt_range,thrust_range,'o')
# plt.show()
print('work = ' + str(work))
## heat transfer required
total_heat_flux = heat_flux(CEA.Pr, CEA.cp, CEA.gamma, CEA.c, CEA.w,
CEA.T_c, T_w, spike)
return (work, total_heat_flux)
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 ideal_thrust(A_t,A_e,gamma,p_c,T_c,rho_c,a_c,altitude,m_dot):
p_atm,T_atm,rho_atm = gd.standard_atmosphere([altitude])
epsilon = A_e/A_t
M_e = optimize.fsolve(lambda M: gd.expansion_ratio_zero(1,M,gamma,epsilon),5)
# print(M_e)
# throat conditions
T_ratio,p_ratio,rho_ratio,a_ratio = gd.isentropic_ratios(0,1,gamma)
T_t = T_ratio*T_c; p_t = p_ratio*p_c; rho_t = rho_ratio*rho_c; a_t = a_ratio*a_c;
# exit conditions
T_ratio,p_ratio,rho_ratio,a_ratio = gd.isentropic_ratios(0,M_e,gamma)
T_e = T_ratio*T_c; p_e = p_ratio*p_c; rho_e = rho_ratio*rho_c; a_e = a_ratio*a_c;
# mass flow rate (constant in ideal case)
# m_dot = rho_t*A_t*a_t*1
# print(m_dot)
# print('rho calc ' + str(rho_t))
# print('a_t calc ' + str(a_t))
# print('V_e calc ' + str(M_e*a_e))
# print('p_e - p_atm' + str(p_e - p_atm))
# print('M_e ' + str(M_e))
return m_dot*M_e*a_e #+ (p_e - p_atm)*A_e
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 cost_func_contour_params(params,
spike,
T_w,
CEA,
alpha,
beta,
chr_mesh_n=145,
no_core=4):
params = np.asarray(params)
if len(params.shape) > 1:
raise ValueError('Input params not correct shape. Should be 1D array')
x_vals, y_vals = np.split(params, 2)
spike.x = x_vals
spike.y = y_vals
## thurst estimation over altitude
alt_range = np.linspace(0, 9144, 3 * no_core)
# shuffle arrays so each core computes similar complexity on average
np.random.shuffle(alt_range)
(p_atm_r, T_atm_r, rho_atm_r) = gd.standard_atmosphere(alt_range)
thrust_range = multicore_thrust_compute(spike,
alt_range,
CEA.gamma,
downstream_factor=1.2,
chr_mesh_n=chr_mesh_n,
no_core=no_core)
# unshuffle arrays
ordered_idx = np.argsort(alt_range)
alt_range = alt_range[ordered_idx]
thrust_range = thrust_range[ordered_idx]
work = np.trapz(alt_range, thrust_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)
return -alpha * work + beta * total_heat_flux
def COST_FNC(params, spike, T_w, CEA, alpha, beta, chr_mesh_n=145, no_core=4):
try:
x_vals, y_vals = params
except ValueError:
x_vals, y_vals = np.split(params, 2)
spike.x = x_vals
spike.y = y_vals
### CALCULATING COST
## thurst estimation over altitude
alt_range = np.linspace(0, 9144, 3 * no_core)
np.random.shuffle(alt_range)
(p_atm_r, T_atm_r, rho_atm_r) = gd.standard_atmosphere(alt_range)
#print(CEA.p_c/p_atm_r)
#thrust_range = multicore_thrust_compute(spike,alt_range,CEA.gamma,downstream_factor=1.2,chr_mesh_n=50,no_core=4)
thrust_range = multicore_thrust_compute(spike,
alt_range,
CEA.gamma,
downstream_factor=1.2,
chr_mesh_n=chr_mesh_n,
no_core=no_core)
# unshuffle of arrays
ordered_idx = np.argsort(alt_range)
alt_range = alt_range[ordered_idx]
thrust_range = thrust_range[ordered_idx]
work = np.trapz(thrust_range, alt_range)
print("work = " + str(work) + ". Bounds [" + str(alt_range[0]) + ', ' +
str(alt_range[-1]) + ']')
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
R = (1 - 1 / gamma) * 1.8292 * 1000 #1.8292 1.6196
a_c = np.sqrt(gamma * (1 - 1 / gamma) * 200.07 * T_c)
n = 1000
spike = plug_nozzle(expansion_ratio,
A_t,
r_e,
gamma,
T_c,
p_c,
a_c,
rho_c,
n,
truncate_ratio=1.0)
alts = np.linspace(0, 15000, 100)
p_atms, T_atms, rho_atms = gd.standard_atmosphere(alts)
thrusts = np.zeros(p_atms.shape)
for i in range(len(p_atms)):
thrusts[i] = spike.calc_ideal_thrust(p_atms[i])
plt.plot(alts, thrusts)
# MOC_mesh = chr_mesh(spike,gamma,0,50,downstream_factor=1.2,plot_chr=1)
#plt.plot(MOC_mesh.x,MOC_mesh.y,'ro')
#plt.plot(MOC_mesh.x[MOC_mesh.ID_jet_boundary],MOC_mesh.y[MOC_mesh.ID_jet_boundary],'go-')
# MOC_mesh.compute_thrust('nearest',10)
#plt.plot(MOC_mesh.x[MOC_mesh.ID_contour_chr],MOC_mesh.y[MOC_mesh.ID_contour_chr],'bo-')
beta,
design_alt_init,
truncate_ratio_init,
chr_mesh_n=120,
no_alt_range=30,
no_core=1)
except:
pass
#design diverging section, 20% truncation
plug1 = opt_aero.spike_init
plug1.define_compression(1.15 / 1000, 4.51 / 1000, 1, 12.91 / 1000, 10000)
print('Thrust = ' +
str(plug1.calc_ideal_thrust(gd.standard_atmosphere([9144])[0])))
print('Throat area = ' + str(plug1.A_t))
print('Expansion ratio = ' + str(plug1.expansion_ratio))
#plug1.plot_contour(plt)
#plt.axis('equal')
tck = interpolate.splrep(plug1.x, plug1.y)
plt.plot(plug1.x, plug1.y, plug1.x, interpolate.splev(plug1.x, tck), 'ro')
init_angle = np.arctan(interpolate.splev(plug1.x[0], tck, der=1))
print("inital angle: " + str(init_angle * 180 / np.pi))
plt.show()
plug1.save_to_csv()
def record_breaker_nozzle():
r_e = 0.027 # r_e = 0.056 #0.034 # likely too large
## CONSTANTS OF DESIGN FOR HEAT FLUX
#user input variable in metric units:
T_w = 600 #[K] desired temperature of nozzle
## CONSTANTS OF SIM
alpha = 1 #0.07/8 # 0.07/8 : 1 ratio of alpha : beta gives very similar weights
beta = 0 #1
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
plug1 = plug_nozzle(design_alt,
r_e,
gamma,
T_c,
p_c,
a_c,
rho_c,
100,
truncate_ratio=truncate_ratio)
# zero_x = np.linspace(plug1.x[-1],plug1.length,20)
# zero_y = np.zeros(zero_x.shape)
(p_atm, T_atm, rho_atm) = gd.standard_atmosphere([design_alt])
print('Ideal thrust: ' + str(plug1.calc_ideal_thrust(p_atm)))
fig, ax1 = plt.subplots(1, 1)
plug1.plot_contour(ax1)
ax1.set_title('Mach Isentropic Expansion Contour Plot')
# m_dot = plug1.rho[0]*plug1.A_t*plug1.V[0]
# print("Mass flow rate: " + str(m_dot))
plug1.plot_exhaust_contourf(ax1, plug1.M, n=100)
ax1.set_xlabel('x (m)')
ax1.set_ylabel('y (m)')
# plug1.save_to_csv()
name = "final_report/mach_isentropic_plot"
fig.set_size_inches(18.5, 10.5)
plt.savefig(name, dpi=100)
plt.show()
plt.close()
# print(plug1.alpha)
## OTHER PLOTS
fig, (ax2, ax3, ax4) = plt.subplots(3, 1)
plug1.plot_contour(ax2)
ax2.set_title('Temperature Isentropic Expansion Contour Plot')
# m_dot = plug1.rho[0]*plug1.A_t*plug1.V[0]
# print("Mass flow rate: " + str(m_dot))
plug1.plot_exhaust_contourf(ax2, plug1.T, n=100)
ax2.set_xlabel('x (m)')
ax2.set_ylabel('y (m)')
# plug1.save_to_csv()
fig.set_size_inches(18.5, 10.5)
####
plug1.plot_contour(ax3)
ax3.set_title('Pressure Isentropic Expansion Contour Plot')
# m_dot = plug1.rho[0]*plug1.A_t*plug1.V[0]
# print("Mass flow rate: " + str(m_dot))
plug1.plot_exhaust_contourf(ax3, plug1.p, n=100)
ax3.set_xlabel('x (m)')
ax3.set_ylabel('y (m)')
# plug1.save_to_csv()
fig.set_size_inches(18.5, 10.5)
plug1.plot_contour(ax4)
ax4.set_title('Density Isentropic Expansion Contour Plot')
# m_dot = plug1.rho[0]*plug1.A_t*plug1.V[0]
# print("Mass flow rate: " + str(m_dot))
plug1.plot_exhaust_contourf(ax4, plug1.rho, n=100)
ax4.set_xlabel('x (m)')
ax4.set_ylabel('y (m)')
# plug1.save_to_csv()
name = "final_report/T_p_rho"
fig.set_size_inches(18.5, 10.5)
plt.savefig(name, dpi=100)
plt.show()
plt.close()
def test_stand_nozzle():
r_e = 0.027 #0.034 # likely too large
## CONSTANTS OF DESIGN FOR HEAT FLUX
#user input variable in metric units:
T_w = 600 #[K] desired temperature of nozzle
## CONSTANTS OF SIM
alpha = 1 #0.07/8 # 0.07/8 : 1 ratio of alpha : beta gives very similar weights
beta = 0 #1
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
plug1 = plug_nozzle(design_alt,
r_e,
gamma,
T_c,
p_c,
a_c,
rho_c,
100,
truncate_ratio=1)
print('Throat area: ' + str(plug1.A_t))
(p_atm, T_atm, rho_atm) = gd.standard_atmosphere([design_alt])
def make_isentropic_plots(plug1, p_atm):
print('Ideal thrust: ' + str(plug1.calc_ideal_thrust(p_atm)))
fig, ax1 = plt.subplots(1, 1)
plug1.plot_contour(ax1)
ax1.set_title('Mach Isentropic Expansion Contour Plot')
# m_dot = plug1.rho[0]*plug1.A_t*plug1.V[0]
# print("Mass flow rate: " + str(m_dot))
plug1.plot_exhaust_contourf(ax1, plug1.M, n=100)
ax1.set_xlabel('x (m)')
ax1.set_ylabel('y (m)')
# plug1.save_to_csv()
name = "final_report/mach_isen_plot_lab"
fig.set_size_inches(18.5, 10.5)
plt.savefig(name, dpi=100)
plt.show()
plt.close()
# print(plug1.alpha)
## OTHER PLOTS
fig, (ax2, ax3, ax4) = plt.subplots(3, 1)
plug1.plot_contour(ax2)
ax2.set_title('Temperature Isentropic Expansion Contour Plot')
# m_dot = plug1.rho[0]*plug1.A_t*plug1.V[0]
# print("Mass flow rate: " + str(m_dot))
plug1.plot_exhaust_contourf(ax2, plug1.T, n=100)
ax2.set_xlabel('x (m)')
ax2.set_ylabel('y (m)')
# plug1.save_to_csv()
fig.set_size_inches(18.5, 10.5)
####
plug1.plot_contour(ax3)
ax3.set_title('Pressure Isentropic Expansion Contour Plot')
# m_dot = plug1.rho[0]*plug1.A_t*plug1.V[0]
# print("Mass flow rate: " + str(m_dot))
plug1.plot_exhaust_contourf(ax3, plug1.p, n=100)
ax3.set_xlabel('x (m)')
ax3.set_ylabel('y (m)')
# plug1.save_to_csv()
fig.set_size_inches(18.5, 10.5)
plug1.plot_contour(ax4)
ax4.set_title('Density Isentropic Expansion Contour Plot')
# m_dot = plug1.rho[0]*plug1.A_t*plug1.V[0]
# print("Mass flow rate: " + str(m_dot))
plug1.plot_exhaust_contourf(ax4, plug1.rho, n=100)
ax4.set_xlabel('x (m)')
ax4.set_ylabel('y (m)')
# plug1.save_to_csv()
name = "final_report/isen_plots_lab"
fig.set_size_inches(18.5, 10.5)
plt.savefig(name, dpi=100)
plt.show()
plt.close()
def make_MOC_plot(plug1, design_alt, gamma, n):
#mesh plot
plug1.plot_contour(plt)
plt.xlabel('x (m)')
plt.ylabel('y (m)')
plt.title('Mesh generated by M.O.C.')
plug_mesh = chr_mesh(plug1,
gamma,
design_alt,
30,
downstream_factor=1.1,
plot_chr=1)
# plug1.save_to_csv()
name = "final_report/MOC_mesh"
plt.axis('equal')
plt.savefig(name, dpi=100)
plt.show()
plt.close()
# mach plot
plug_mesh = chr_mesh(plug1,
gamma,
design_alt,
120,
downstream_factor=1.0,
plot_chr=0)
thrust = plug_mesh.compute_thrust('linear', 100)
print('MOC thrust: ' + str(thrust))
tck = interpolate.splrep(plug_mesh.x[plug_mesh.ID_contour_chr],
plug_mesh.y[plug_mesh.ID_contour_chr])
fig, (ax1) = plt.subplots(1, 1)
plug1.plot_contour(ax1)
ax1.axes.get_xaxis().set_ticklabels([])
ax1.axes.get_yaxis().set_ticklabels([])
X_plt = np.linspace(0, plug_mesh.x.max(), 100)
Y_plt = np.linspace(0, plug_mesh.y.min(), 100)
X_plt, Y_plt = np.meshgrid(X_plt, Y_plt)
invalid_grid = interpolate.splev(X_plt.flatten(),
tck) < Y_plt.flatten()
invalid_grid = invalid_grid.reshape(X_plt.shape)
M_contour = interpolate.griddata((plug_mesh.x, plug_mesh.y),
plug_mesh.M, (X_plt, Y_plt),
method='linear')
#ax1.plot(X_plt[valid_grid],Y_plt[invalid_grid],'.')
M_contour[invalid_grid] = np.nan
# ax1.plot(X_plt.flatten(),interpolate.splev(X_plt.flatten(),tck),'.')
ax1.set_aspect('equal', 'box')
for item in ([ax1.title, ax1.xaxis.label, ax1.yaxis.label] +
ax1.get_xticklabels() + ax1.get_yticklabels()):
item.set_fontsize(20)
M_fill = ax1.contourf(X_plt, -Y_plt, M_contour, cmap=cm.jet)
cbar = plt.colorbar(M_fill, ax=ax1, orientation='horizontal')
cbar.ax.tick_params(labelsize=20)
name = "final_report/M_contour_MOC"
fig.set_size_inches(18.5, 10.5)
# ax1.set_xlabel('x (m)')
# ax1.set_ylabel('y (m)')
# ax1.set_title('M Contour Plot for M.O.C.')
ax1.set_aspect('equal', 'box')
plt.savefig(name, dpi=100)
plt.show()
plt.close()
fig, (ax2) = plt.subplots(1, 1)
ax2.axes.get_xaxis().set_ticklabels([])
ax2.axes.get_yaxis().set_ticklabels([])
for item in ([ax2.title, ax2.xaxis.label, ax2.yaxis.label] +
ax2.get_xticklabels() + ax2.get_yticklabels()):
item.set_fontsize(20)
# velocity plot
U = np.cos(plug_mesh.theta) * plug_mesh.V
V = np.sin(plug_mesh.theta) * plug_mesh.V
x_grid, y_grid = np.meshgrid(np.linspace(0, plug_mesh.x.max(), 30),
np.linspace(0, plug_mesh.y.min(), 20))
invalid_grid = interpolate.splev(x_grid.flatten(),
tck) < y_grid.flatten()
invalid_grid = invalid_grid.reshape(x_grid.shape)
u_grid = interpolate.griddata((plug_mesh.x, plug_mesh.y),
U, (x_grid, y_grid),
method='linear')
v_grid = interpolate.griddata((plug_mesh.x, plug_mesh.y),
V, (x_grid, y_grid),
method='linear')
mag_grid = interpolate.griddata((plug_mesh.x, plug_mesh.y),
plug_mesh.V, (x_grid, y_grid),
method='linear')
mag_grid[invalid_grid] = np.nan
plug1.plot_contour(ax2)
fig.set_size_inches(18.5, 10.5)
Q = ax2.quiver(x_grid,
-y_grid,
u_grid,
-v_grid,
mag_grid,
cmap=cm.jet,
alpha=0.5)
cbar = plt.colorbar(Q, ax=ax2, orientation="horizontal")
cbar.ax.tick_params(labelsize=20)
ax2.set_aspect('equal', 'box')
# ax2.set_title('Velocity Vector Plot (m/s)')
# ax2.set_xlabel('x (m)')
# ax2.set_ylabel('y (m)')
name = "final_report/V_vec"
plt.savefig(name, dpi=100)
plt.show()
# other property contours
make_MOC_plot(plug1, design_alt, gamma, 10)
def __init__(self,spike,gamma,altitude,n,downstream_factor=1.1,plot_chr=0,clean_mesh=1):
self.spike =copy.copy(spike); self.gamma = gamma; self.altitude =altitude; self.n = n
self.downstream_factor = downstream_factor # percentage down after mesh cross with centre to continue meshing
# constants of iteration
self.plot_chr = plot_chr
self.flip_plug() # flips sign of plug if required
self.slope_init = np.arctan(-(self.spike.lip_x-self.spike.x[0])/(self.spike.lip_y-self.spike.y[0]));
#print(self.slope_init*180/np.pi)
self.tck = interpolate.splrep(self.spike.x,self.spike.y,full_output=0)
(self.p_atm,self.T_atm,self.rho_atm) = gd.standard_atmosphere([altitude])
self.gamma = gamma
self.PR = self.spike.p_c/(self.p_atm)
self.V_l = np.sqrt(2/(self.gamma-1))*self.spike.a_c
# Point storage
self.chr_array = np.array([]) # array for storing points based on ID
# logs characteristics that have not been intercepted
self.ID_left_chr = []
self.ID_right_chr = []
self.ID_jet_boundary = []
self.ID_contour_chr = []
#self.ID_next_chr_jet = []
#self.ID_compression_chr = []
self.ID = 0 # ID of next point, not yet computed
# COMPUTE EXPANSION FAN POINTS (initial data line)
# construction of initial expansion fan
self.compute_initial_expansion_fan()
# TODO: COMPUTE CHAR. NET DOWN TO NOZZLE BASE (when A.x > spike.x[-1], switch centre point jet boundary)
## after computing initial fan, load all point id's into ordered list showing right and left running characteristics that have not been used to compute a downstream point
# TODO: COMPUTE NOZZLE BASE PRESSURE
# while (self.chr_array[self.ID_contour_chr[0]].x <= spike.x.max()):
# pass
# TODO: COMPUTE PLUME DOWNSTREAM OF BASE
# TODO: COMPUTE THRUST PRODUCED
# for point in self.chr_array:
# plt.plot(point.x,point.y,'rX')
# print(self.ID_left_chr)
# print(self.ID_right_chr)
# print(self.ID_jet_boundary)
# print(self.ID_contour_chr)
# base conditions
self.p_b = self.base_pressure()
self.contour_fan = 1
self.new_fan = 1
self.first_base_intercept = 1
self.centre_line_intercept = 0
self.END_SIM = 0
#while (self.chr_array[self.ID_contour_chr[-1]].x <= spike.x.max()):
while self.chr_point_less_zero() and self.contour_converge():
## TODO: COMPUTE EXPANSION FAN UNTIL POINT IS > spike.length in which case, remove from all tracking lists and do not add to chr_array
## CONTOUR FAN
# if (self.contour_fan):
ID_temp = self.ID_right_chr.pop(0)
#print(ID_temp)
if(self.new_fan):
# intersection
self.new_fan = 0
if (self.on_nozzle_contour(self.chr_array[ID_temp])):
new_point = self.contour_point(self.chr_array[ID_temp],plot_chr=self.plot_chr)
self.chr_array = np.append(self.chr_array,new_point)
self.ID += 1
# first point
ID_temp = self.ID_right_chr.pop(0)
new_point = self.general_point(self.chr_array[ID_temp],self.chr_array[self.ID-1],plot_chr=self.plot_chr)
self.chr_array = np.append(self.chr_array,new_point)
self.ID += 1
else:
if (self.first_base_intercept):
self.first_base_intercept = 0
first_base_point = self.chr_array[self.ID_contour_chr[-1]]
#print(self.spike.p_c/self.p_b)
M_b = gd.PR_expansion_mach(self.spike.p_c/self.p_b,self.gamma)
#print(M_b)
theta_b = gd.prandtl_meyer(M_b,self.gamma) - gd.prandtl_meyer(first_base_point.M,self.gamma)
W_b = first_base_point.W
first_base_point = chr_point(self.gamma,self.spike.x[-1],self.spike.y[-1],theta_b,W_b,'contour')
new_point = self.internal_jet_point(first_base_point,self.chr_array[ID_temp],plot_chr=self.plot_chr)
self.chr_array = np.append(self.chr_array,new_point)
self.ID += 1
else:
new_point = self.internal_jet_point(self.chr_array[self.ID_contour_chr[-1]],self.chr_array[ID_temp],plot_chr=self.plot_chr)
self.chr_array = np.append(self.chr_array,new_point)
self.ID += 1
elif(ID_temp==-1):
# self.ID_next_chr_jet.append(self.ID_left_chr.pop(0))
# plt.plot(self.chr_array[self.ID_jet_boundary[-1]].x,self.chr_array[self.ID_jet_boundary[-1]].y,'gx')
new_point = self.general_point(self.chr_array[self.ID_jet_boundary[-1]],self.chr_array[self.ID-1],plot_chr=self.plot_chr)
self.chr_array = np.append(self.chr_array,new_point)
#self.ID_left_chr.append(self.ID)
self.ID += 1
new_point = self.jet_boundary_point(self.chr_array[self.ID_jet_boundary[-1]],self.chr_array[self.ID-1],plot_chr=self.plot_chr)
self.chr_array = np.append(self.chr_array,new_point)
#plt.plot(new_point.x,new_point.y,'gx')
#self.ID_jet_boundary.append(self.ID)
self.ID +=1
self.contour_fan = 0
self.add_break_ID()
self.new_fan = 1
# if (self.centre_line_intercept):
# self.END_SIM = 1
else:
if(self.chr_array[ID_temp].pt_type!="same_fam"):
temp1 = self.same_fam_point(self.chr_array[self.ID_right_chr[0]],self.chr_array[ID_temp])
temp2 = self.general_point(self.chr_array[ID_temp],self.chr_array[self.ID-1])
if (temp1.x < temp2.x) and (temp1.x>self.chr_array[ID_temp].x) :
#self.ID_right_chr.pop(-1); self.ID_left_chr.pop(-1)
#self.compression_offset += 1
#self.plot_chr=1
self.ID_left_chr.pop(-1)
self.ID_right_chr.pop(-1)
#print(temp1.x)
if (self.plot_chr):
plt.plot(self.chr_array[self.ID_right_chr[0]].x,self.chr_array[self.ID_right_chr[0]].y,'bx',self.chr_array[ID_temp].x,self.chr_array[ID_temp].y,'rx')
plt.plot(temp1.x,temp1.y,'go')
#plt.plot([self.chr_array[ID_temp].x, temp1.x],[self.chr_array[ID_temp].y, temp1.y])
self.chr_array = np.append(self.chr_array,temp1)
new_point = self.general_point(self.chr_array[-1],self.chr_array[self.ID-1],plot_chr=self.plot_chr)
#plt.plot(self.chr_array[self.ID-1].x,self.chr_array[self.ID-1].y,'ro')
self.ID += 1
else:
new_point = temp2
if (new_point.x<=self.spike.length):
pass
if (self.plot_chr):
plt.plot([self.chr_array[ID_temp].x, temp2.x],[self.chr_array[ID_temp].y, temp2.y],'k',[self.chr_array[self.ID-1].x, temp2.x],[self.chr_array[self.ID-1].y, temp2.y],'k')
self.ID += 1
self.chr_array = np.append(self.chr_array,new_point)
# ## JET BOUNDARY FAN
# else:
# ID_temp = self.ID_right_chr.pop(0)
# if(self.new_fan):
# new_point = self.general_point(self.chr_array[ID_temp],self.chr_array[self.ID_contour_chr[-1]])
# self.chr_array = np.append(self.chr_array,new_point)
# self.ID += 1
# self.new_fan = 0
# elif (ID_temp == -1):
# new_point=self.jet_boundary_point(self.chr_array[self.ID_next_chr_jet[-1]],self.chr_array[ID_temp -1])
# self.chr_array = np.append(self.chr_array, new_point)
# self.ID += 1
# self.contour_fan = 1
# self.new_fan = 1
# self.add_break_ID()
# else:
# new_point = self.general_point(self.chr_array[ID_temp],self.chr_array[ID_temp-1])
# self.chr_array = np.append(self.chr_array,new_point)
# self.ID += 1
## END OF MOC SECTION #####
# function order is important
if(clean_mesh):
self.clean_data()
self.to_arrays()
self.calc_flow_properties()
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
if __name__ == '__main__': # engine constants r_e = 0.027 A_e = np.pi*r_e**2 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 altitude = 9144 p_atm,T_atm,rho_atm = gd.standard_atmosphere([altitude]) plug1 = plug_nozzle(altitude,r_e,gamma,T_c,p_c,a_c,rho_c,100,truncate_ratio = 1) m = (plug1.lip_y - plug1.y[0])/(plug1.lip_x - plug1.x[0]) # simulation constant num_t = 10 t_vec = np.linspace(0,4,num_t) alpha = 0.03/1000 beta = 0.05/1000 thrust_vector_alpha = ideal_thrust_vector(alpha,t_vec,plug1,A_e,gamma,p_c,T_c,rho_c,a_c,altitude)
#user input variable in metric units:
T_w = 600 #[K] desired temperature of nozzle
## CONSTANTS OF SIM
alpha = 1 #0.07/8 # 0.07/8 : 1 ratio of alpha : beta gives very similar weights
beta = 0 #1
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,
