mdot = 4.757 #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
'''Engine geometry'''
Ac = np.pi * 0.1**2 #Chamber cross-sectional area (m^2)
L_star = 1.5 #L_star = Volume_c/Area_t
inner_wall_thickness = 2e-3
outer_wall_thickness = 4e-3
'''Coolant jacket'''
inner_wall_material = bam.materials.CopperC700
outer_wall_material = bam.materials.StainlessSteel304
mdot_coolant = mdot / (OF_ratio + 1)
inlet_T = 298.15 # Coolant inlet temperature
thermo_coolant = thermo.chemical.Chemical('isopropanol')
coolant_transport = cool.TransportProperties(model="thermo",
thermo_object=thermo_coolant,
force_phase='l')
'''Create the engine object'''
perfect_gas = bam.PerfectGas(gamma=gamma, molecular_weight=molecular_weight)
chamber = bam.ChamberConditions(pc, Tc, mdot)
nozzle = bam.Nozzle.from_engine_components(perfect_gas,
chamber,
p_amb,
type="rao",
length_fraction=0.8)
engine = bam.Engine(perfect_gas, chamber, nozzle)
chamber_length = L_star * nozzle.At / Ac
'''Choose the models we want to use for transport properties of the exhaust gas'''
thermo_gas = thermo.mixture.Mixture(
['N2', 'H2O', 'CO2'],
ws=[0.49471, 0.14916, 0.07844
e.add_propellants_by_mass([(ipa, 1-water_mass_fraction),
(water, water_mass_fraction),
(n2o, OF_ratio)])
#Adiabatic combustion using chamber pressure
e.set_state(P = pc/1e5, type='HP')
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(['Isopropyl Alcohol', '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)
coolant_transport = cool.TransportProperties(model = "CoolProp", coolprop_name = f"ETHANOL[{1 - water_mass_fraction}]&WATER[{water_mass_fraction}]")
'''Create the engine object'''
perfect_gas = bam.PerfectGas(gamma = gamma, cp = cp) #Gas for frozen flow
chamber = bam.ChamberConditions(pc, Tc, mdot)
nozzle = bam.Nozzle.from_engine_components(perfect_gas, chamber, p_amb, type = "rao", length_fraction = 0.8)
white_dwarf = bam.Engine(perfect_gas, chamber, nozzle)
chamber_length = L_star*nozzle.At/Ac
print(f"Sea level thrust = {white_dwarf.thrust(1e5)/1000} kN")
print(f"Sea level Isp = {white_dwarf.isp(1e5)} s")
'''Cooling system setup'''
white_dwarf.add_geometry(chamber_length, Ac, inner_wall_thickness, outer_wall_thickness)
white_dwarf.add_exhaust_transport(gas_transport)
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), (n2o, OF_ratio)])
#Adiabatic combustion using chamber pressure
e.set_state(P=pc / 1e5, type='HP')
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 exhaust gas'''
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)
'''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)
white_dwarf.add_ablative(bam.materials.Graphite,
bam.materials.CopperC700,
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)