def calculate(self, prop, p1, pe):
try:
self.calc_prop = prop
self.p1 = p1/14.696 #converting psi to atm
self.pe = pe/14.696
ppp.init()
self.p = ppp.FrozenPerformance()
self.calc_prop_var = [0]*len(self.calc_prop)
list_tuple = []
self.prop_weight = 0
self.propellant_list = []
for i in range(0, len(self.calc_prop)):
self.propellant_list.append(self.calc_prop[i][0])
self.calc_prop_var[i] = ppp.PROPELLANTS[self.calc_prop[i][0]]
self.prop_weight += float(self.calc_prop[i][1])
list_tuple.append((self.calc_prop_var[i], float(self.calc_prop[i][1])))
self.p.add_propellants_by_mass(list_tuple)
self.p.set_state(P = float(self.p1), Pe = float(self.pe))
self.calculate_button.config(state = 'disabled')
self.display_button.config(state = 'active')
self.reset_button.grid(column = 2, row = 7, columnspan = 2, rowspan = 2, sticky = 'n')
self.validation_label.configure(text = "Total Weight: " + str(self.prop_weight), fg = 'white', bg = "#222831")
#self.validation_label.configure(text = "Sucess :)", bg = 'green', fg = 'white', font = ("Garamond", 20))
propellant_calculations.prop_variable = self.p
global outsidep
outsidep = self.p
except Exception as e:
print("ERROR")
self.validation_label.configure(text = "System Error :/", fg = 'white', bg = 'red')
import numpy as np
from matplotlib import pyplot as plt
from matplotlib import rc
from matplotlib.ticker import FormatStrFormatter
import pandas as pd
import pypropep as ppp
ppp.init()
rc('figure', figsize=(4, 2.5))
rc('legend', fontsize='small')
rc('font', family='serif')
rc('xtick', labelsize='small')
rc('ytick', labelsize='small')
with open('JANAF_O2_g.txt', 'r') as f:
o2_tbl = pd.read_csv(f, sep='\t', header=1, na_values='INFINITE')
with open('JANAF_O_g.txt', 'r') as f:
o_tbl = pd.read_csv(f, sep='\t', header=1, na_values='INFINITE')
def f_T(tbl, T, name):
return np.interp(T, tbl['T(K)'], tbl[name])
R = 8.314
Tref = 298.15
Pref = 1e5
Rref = Pref / R / Tref
T = np.linspace(500, 6000, 30)
def pypropep():
import pypropep
pypropep.init()
return pypropep
def test_thermo_file_override(pypropep):
therm_file = os.path.dirname(pypropep.__file__) + '/data/thermo.dat'
pypropep.init(thermo_file=therm_file)
def test_prop_file_override(pypropep):
prop_file = os.path.dirname(pypropep.__file__) + '/data/propellant.dat'
pypropep.init(propellant_file=prop_file)
def white_dwarf_cooling_data(water_mass_fraction):
'''Engine dimensions'''
Ac = np.pi*0.1**2 #Chamber cross-sectional area (m^2)
L_star = 1.5 #L_star = Volume_c/Area_t
wall_thickness = 2e-3
'''Chamber conditions'''
pc = 15e5 #Chamber pressure (Pa)
p_tank = 20e5 #Tank / inlet coolant stagnation pressure (Pa) - used for cooling jacket
mdot = 5.4489 #Mass flow rate (kg/s)
p_amb = 1.01325e5 #Ambient pressure (Pa). 1.01325e5 is sea level atmospheric.
OF_ratio = 3.5 #Oxidiser/fuel mass ratio
'''Coolant jacket'''
wall_material = bam.materials.CopperC700
mdot_coolant = mdot/(OF_ratio + 1)
inlet_T = 298.15 #Coolant inlet temperature
'''Get combustion properties from pypropep'''
#Initialise and get propellants
ppp.init()
e = ppp.Equilibrium()
p = ppp.ShiftingPerformance()
ipa = ppp.PROPELLANTS['ISOPROPYL ALCOHOL']
water = ppp.PROPELLANTS['WATER']
n2o = ppp.PROPELLANTS['NITROUS OXIDE']
#Add propellants by mass fractions (note the mass fractions can add up to more than 1)
e.add_propellants_by_mass([(ipa, 1),
(water, water_mass_fraction),
(n2o, OF_ratio)])
p.add_propellants_by_mass([(ipa, 1),
(water, water_mass_fraction),
(n2o, OF_ratio)])
#Adiabatic combustion
e.set_state(P = pc/p_amb, type = 'HP')
#Set chamber pressure and exit pressure in atmospheres
p.set_state(P = pc/p_amb, Pe = 1)
gamma = e.properties.Isex #I don't know why they use 'Isex' for gamma.
cp = 1000*e.properties.Cp #Cp is given in kJ/kg/K, we want J/kg/K
Tc = e.properties.T
'''Choose the models we want to use for transport properties of the coolant and exhaust gas'''
thermo_coolant = thermo.mixture.Mixture(['isopropanol', 'water'], ws = [1 - water_mass_fraction, water_mass_fraction])
thermo_gas = thermo.mixture.Mixture(['N2', 'H2O', 'CO2'], zs = [e.composition['N2'], e.composition['H2O'], e.composition['CO2']])
gas_transport = cool.TransportProperties(model = "thermo", thermo_object = thermo_gas, force_phase = 'g')
coolant_transport = cool.TransportProperties(model = "thermo", thermo_object = thermo_coolant, force_phase = 'l')
'''Create the engine object'''
perfect_gas = bam.PerfectGas(gamma = gamma, cp = cp) #Gas for frozen flow
chamber_conditions = bam.ChamberConditions(pc, Tc, mdot)
nozzle = bam.Nozzle.from_engine_components(perfect_gas, chamber_conditions, p_amb, type = "rao", length_fraction = 0.8)
white_dwarf = bam.Engine(perfect_gas, chamber_conditions, nozzle)
chamber_length = L_star*nozzle.At/Ac
'''Add the cooling system to the engine'''
white_dwarf.add_geometry(chamber_length, Ac, wall_thickness)
white_dwarf.add_exhaust_transport(gas_transport)
#Spiral channels
#white_dwarf.add_cooling_jacket(wall_material, inlet_T, p_tank, coolant_transport, mdot_coolant,
# configuration = "spiral", channel_shape = "semi-circle", channel_width = 0.020)
#Or vertical channels
white_dwarf.add_cooling_jacket(wall_material, inlet_T, p_tank, coolant_transport, mdot_coolant, configuration = "vertical", channel_height = 0.001)
'''Run the heating analysis'''
cooling_data = white_dwarf.steady_heating_analysis(number_of_points = 250, to_json = False)
return cooling_data, white_dwarf.isp(p_amb)