PhlyGreen Tutorial: sizing a parallel-hybrid aircraftĀ¶
Import packages
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import sys
sys.path.insert(0,'../')
import PhlyGreen as pg
import numpy as np
import matplotlib.pyplot as plt
import sys sys.path.insert(0,'../') import PhlyGreen as pg import numpy as np import matplotlib.pyplot as plt
Create instances of subsystems. Each one of those requires an Aircraft object as argument. By passing (None) as argument, we are creating empty objects (for the moment).
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powertrain = pg.Systems.Powertrain.Powertrain(None)
structures = pg.Systems.Structures.Structures(None)
aerodynamics = pg.Systems.Aerodynamics.Aerodynamics(None)
performance = pg.Performance.Performance(None)
mission = pg.Mission.Mission(None)
weight = pg.Weight.Weight(None)
constraint = pg.Constraint.Constraint(None)
welltowake = pg.WellToWake.WellToWake(None)
battery = pg.Systems.Battery.Battery(None)
climateimpact = pg.ClimateImpact.ClimateImpact(None)
powertrain = pg.Systems.Powertrain.Powertrain(None) structures = pg.Systems.Structures.Structures(None) aerodynamics = pg.Systems.Aerodynamics.Aerodynamics(None) performance = pg.Performance.Performance(None) mission = pg.Mission.Mission(None) weight = pg.Weight.Weight(None) constraint = pg.Constraint.Constraint(None) welltowake = pg.WellToWake.WellToWake(None) battery = pg.Systems.Battery.Battery(None) climateimpact = pg.ClimateImpact.ClimateImpact(None)
Create instance of Aircraft object.
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myaircraft = pg.Aircraft(powertrain, structures, aerodynamics, performance, mission, weight, constraint, welltowake, battery, climateimpact)
myaircraft = pg.Aircraft(powertrain, structures, aerodynamics, performance, mission, weight, constraint, welltowake, battery, climateimpact)
Create the connections with its subsystems.
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powertrain.aircraft = myaircraft
structures.aircraft = myaircraft
aerodynamics.aircraft = myaircraft
mission.aircraft = myaircraft
performance.aircraft = myaircraft
weight.aircraft = myaircraft
constraint.aircraft = myaircraft
welltowake.aircraft = myaircraft
battery.aircraft = myaircraft
climateimpact.aircraft = myaircraft
powertrain.aircraft = myaircraft structures.aircraft = myaircraft aerodynamics.aircraft = myaircraft mission.aircraft = myaircraft performance.aircraft = myaircraft weight.aircraft = myaircraft constraint.aircraft = myaircraft welltowake.aircraft = myaircraft battery.aircraft = myaircraft climateimpact.aircraft = myaircraft
Define input dictionaries
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ConstraintsInput = {'DISA': 0.,
'Cruise': {'Speed': 0.5, 'Speed Type':'Mach', 'Beta': 0.95, 'Altitude': 8000.},
'AEO Climb': {'Speed': 210, 'Speed Type':'KCAS', 'Beta': 0.97, 'Altitude': 6000., 'ROC': 5},
'OEI Climb': {'Speed': 1.2*34.5, 'Speed Type': 'TAS', 'Beta': 1., 'Altitude': 0., 'Climb Gradient': 0.021},
'Take Off': {'Speed': 90, 'Speed Type': 'TAS', 'Beta': 1., 'Altitude': 100., 'kTO': 1.2, 'sTO': 950},
'Landing':{'Speed': 59., 'Speed Type': 'TAS', 'Altitude': 500.},
'Turn':{'Speed': 210, 'Speed Type': 'KCAS', 'Beta': 0.9, 'Altitude': 5000, 'Load Factor': 1.1},
'Ceiling':{'Speed': 0.5, 'Beta': 0.8, 'Altitude': 9500, 'HT': 0.5},
'Acceleration':{'Mach 1': 0.3, 'Mach 2':0.4, 'DT': 180, 'Altitude': 6000, 'Beta': 0.9}}
ConstraintsInput = {'DISA': 0., 'Cruise': {'Speed': 0.5, 'Speed Type':'Mach', 'Beta': 0.95, 'Altitude': 8000.}, 'AEO Climb': {'Speed': 210, 'Speed Type':'KCAS', 'Beta': 0.97, 'Altitude': 6000., 'ROC': 5}, 'OEI Climb': {'Speed': 1.2*34.5, 'Speed Type': 'TAS', 'Beta': 1., 'Altitude': 0., 'Climb Gradient': 0.021}, 'Take Off': {'Speed': 90, 'Speed Type': 'TAS', 'Beta': 1., 'Altitude': 100., 'kTO': 1.2, 'sTO': 950}, 'Landing':{'Speed': 59., 'Speed Type': 'TAS', 'Altitude': 500.}, 'Turn':{'Speed': 210, 'Speed Type': 'KCAS', 'Beta': 0.9, 'Altitude': 5000, 'Load Factor': 1.1}, 'Ceiling':{'Speed': 0.5, 'Beta': 0.8, 'Altitude': 9500, 'HT': 0.5}, 'Acceleration':{'Mach 1': 0.3, 'Mach 2':0.4, 'DT': 180, 'Altitude': 6000, 'Beta': 0.9}}
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MissionInput = {'Range Mission': 750, #nautical miles
'Range Diversion': 220, #nautical miles
'Beta start': 0.97,
'Payload Weight': 4560, # [kg]
'Crew Weight': 500} # [kg]
MissionInput = {'Range Mission': 750, #nautical miles 'Range Diversion': 220, #nautical miles 'Beta start': 0.97, 'Payload Weight': 4560, # [kg] 'Crew Weight': 500} # [kg]
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MissionStages = {'Takeoff': {'Supplied Power Ratio':{'phi': 0.}},
'Climb1': {'type': 'ConstantRateClimb', 'input': {'CB': 0.16, 'Speed': 77, 'StartAltitude': 100, 'EndAltitude': 1500}, 'Supplied Power Ratio':{'phi_start': 0, 'phi_end':0}},
'Climb2': {'type': 'ConstantRateClimb', 'input': {'CB': 0.08, 'Speed': 120, 'StartAltitude': 1500, 'EndAltitude': 4500}, 'Supplied Power Ratio':{'phi_start': 0, 'phi_end':0 }},
'Climb3': {'type': 'ConstantRateClimb', 'input': {'CB': 0.07, 'Speed': 125, 'StartAltitude': 4500, 'EndAltitude': 8000}, 'Supplied Power Ratio':{'phi_start': 0, 'phi_end':0 }},
'Cruise': {'type': 'ConstantMachCruise', 'input':{ 'Mach': 0.4, 'Altitude': 8000}, 'Supplied Power Ratio':{'phi_start': 0, 'phi_end':0.5}},
'Descent1': {'type': 'ConstantRateDescent', 'input':{'CB': -0.04, 'Speed': 90, 'StartAltitude': 8000, 'EndAltitude': 200}, 'Supplied Power Ratio':{'phi_start': 0.0, 'phi_end':0.0 }}}
DiversionStages = {'Climb1': {'type': 'ConstantRateClimb', 'input': {'CB': 0.08, 'Speed': 110, 'StartAltitude': 200, 'EndAltitude': 3100}, 'Supplied Power Ratio':{'phi_start': 0.0, 'phi_end':0.0 }},
'Cruise': {'type': 'ConstantMachCruise', 'input':{ 'Mach': 0.35, 'Altitude': 3100}, 'Supplied Power Ratio':{'phi_start': 0.0, 'phi_end':0.0}},
'Descent1': {'type': 'ConstantRateDescent', 'input':{'CB': -0.04, 'Speed': 90, 'StartAltitude': 3100, 'EndAltitude': 200}, 'Supplied Power Ratio':{'phi_start': 0.0, 'phi_end':0.0 }}}
MissionStages = {'Takeoff': {'Supplied Power Ratio':{'phi': 0.}}, 'Climb1': {'type': 'ConstantRateClimb', 'input': {'CB': 0.16, 'Speed': 77, 'StartAltitude': 100, 'EndAltitude': 1500}, 'Supplied Power Ratio':{'phi_start': 0, 'phi_end':0}}, 'Climb2': {'type': 'ConstantRateClimb', 'input': {'CB': 0.08, 'Speed': 120, 'StartAltitude': 1500, 'EndAltitude': 4500}, 'Supplied Power Ratio':{'phi_start': 0, 'phi_end':0 }}, 'Climb3': {'type': 'ConstantRateClimb', 'input': {'CB': 0.07, 'Speed': 125, 'StartAltitude': 4500, 'EndAltitude': 8000}, 'Supplied Power Ratio':{'phi_start': 0, 'phi_end':0 }}, 'Cruise': {'type': 'ConstantMachCruise', 'input':{ 'Mach': 0.4, 'Altitude': 8000}, 'Supplied Power Ratio':{'phi_start': 0, 'phi_end':0.5}}, 'Descent1': {'type': 'ConstantRateDescent', 'input':{'CB': -0.04, 'Speed': 90, 'StartAltitude': 8000, 'EndAltitude': 200}, 'Supplied Power Ratio':{'phi_start': 0.0, 'phi_end':0.0 }}} DiversionStages = {'Climb1': {'type': 'ConstantRateClimb', 'input': {'CB': 0.08, 'Speed': 110, 'StartAltitude': 200, 'EndAltitude': 3100}, 'Supplied Power Ratio':{'phi_start': 0.0, 'phi_end':0.0 }}, 'Cruise': {'type': 'ConstantMachCruise', 'input':{ 'Mach': 0.35, 'Altitude': 3100}, 'Supplied Power Ratio':{'phi_start': 0.0, 'phi_end':0.0}}, 'Descent1': {'type': 'ConstantRateDescent', 'input':{'CB': -0.04, 'Speed': 90, 'StartAltitude': 3100, 'EndAltitude': 200}, 'Supplied Power Ratio':{'phi_start': 0.0, 'phi_end':0.0 }}}
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EnergyInput = {'Ef': 43.5*10**6, # [J/kg]
'Contingency Fuel': 130, # [kg]
'Eta Gas Turbine Model': 'constant',
'Eta Gas Turbine': 0.22,
'Eta Gearbox': 0.96,
'Eta Propulsive Model': 'constant',
'Eta Propulsive': 0.9,
'Eta Electric Motor 1': 0.96, #for serial config
'Eta Electric Motor 2': 0.96, #for serial config
'Eta Electric Motor': 0.98, #for parallel config
'Eta PMAD': 0.99,
'Specific Power Powertrain': [3900,7700], # [W/kg]
'Specific Power PMAD': [2200,2200,2200] # [W/kg]
}
EnergyInput = {'Ef': 43.5*10**6, # [J/kg] 'Contingency Fuel': 130, # [kg] 'Eta Gas Turbine Model': 'constant', 'Eta Gas Turbine': 0.22, 'Eta Gearbox': 0.96, 'Eta Propulsive Model': 'constant', 'Eta Propulsive': 0.9, 'Eta Electric Motor 1': 0.96, #for serial config 'Eta Electric Motor 2': 0.96, #for serial config 'Eta Electric Motor': 0.98, #for parallel config 'Eta PMAD': 0.99, 'Specific Power Powertrain': [3900,7700], # [W/kg] 'Specific Power PMAD': [2200,2200,2200] # [W/kg] }
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CellInput = {
'Class': "II",
'Model':'Finger-Cell-Thermal', # used only if Class=II
'SpecificPower': 8000, # [W/kg] if Class=II it can be (i) assigned, or (ii) None and computed from cell properties
'SpecificEnergy': 1500, # [Wh/kg]
'Minimum SOC': 0.2,
'Pack Voltage':800, # [V] used only if Class=II
'Initial temperature': 25, # [Ā°C] used only if Class=II
'Max operative temperature':50 # [Ā°C] used only if Class=II
}
CellInput = { 'Class': "II", 'Model':'Finger-Cell-Thermal', # used only if Class=II 'SpecificPower': 8000, # [W/kg] if Class=II it can be (i) assigned, or (ii) None and computed from cell properties 'SpecificEnergy': 1500, # [Wh/kg] 'Minimum SOC': 0.2, 'Pack Voltage':800, # [V] used only if Class=II 'Initial temperature': 25, # [Ā°C] used only if Class=II 'Max operative temperature':50 # [Ā°C] used only if Class=II }
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AerodynamicsInput = {'AnalyticPolar': {'type': 'Quadratic', 'input': {'AR': 11, 'e_osw': 0.8}},
'Take Off Cl': 1.9,
'Landing Cl': 1.9,
'Minimum Cl': 0.20,
'Cd0': 0.017}
AerodynamicsInput = {'AnalyticPolar': {'type': 'Quadratic', 'input': {'AR': 11, 'e_osw': 0.8}}, 'Take Off Cl': 1.9, 'Landing Cl': 1.9, 'Minimum Cl': 0.20, 'Cd0': 0.017}
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WellToTankInput = {'Eta Charge': 0.95,
'Eta Grid': 1.,
'Eta Extraction': 1.,
'Eta Production': 1.,
'Eta Transportation': 0.25}
ClimateImpactInput = {'H': 100, 'N':1.6e7, 'Y':30, 'EINOx_model':'Filippone', 'WTW_CO2': 8.30e-3, 'Grid_CO2': 9.36e-2}
WellToTankInput = {'Eta Charge': 0.95, 'Eta Grid': 1., 'Eta Extraction': 1., 'Eta Production': 1., 'Eta Transportation': 0.25} ClimateImpactInput = {'H': 100, 'N':1.6e7, 'Y':30, 'EINOx_model':'Filippone', 'WTW_CO2': 8.30e-3, 'Grid_CO2': 9.36e-2}
Set aircraft power configuration and structural weight model
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myaircraft.Configuration = 'Hybrid'
myaircraft.HybridType = 'Parallel'
myaircraft.weight.Class = 'I'
myaircraft.AircraftType = 'ATR'
myaircraft.Configuration = 'Hybrid' myaircraft.HybridType = 'Parallel' myaircraft.weight.Class = 'I' myaircraft.AircraftType = 'ATR'
Import input dictionaries
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myaircraft.ReadInput(AerodynamicsInput,
ConstraintsInput,
MissionInput,
EnergyInput,
MissionStages,
DiversionStages,
LoiterStages=None,
WellToTankInput=WellToTankInput,
CellInput=CellInput,
ClimateImpactInput=ClimateImpactInput,
PropellerInput=None
)
myaircraft.ReadInput(AerodynamicsInput, ConstraintsInput, MissionInput, EnergyInput, MissionStages, DiversionStages, LoiterStages=None, WellToTankInput=WellToTankInput, CellInput=CellInput, ClimateImpactInput=ClimateImpactInput, PropellerInput=None )
Either run the Constraint Analysis....
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myaircraft.constraint.FindDesignPoint()
print('----------------------------------------------')
print(f'Design wing loading W/S: {myaircraft.DesignWTOoS:.1f} [N/m^2]')
print(f'Design power-to-mass ratio P/W: {myaircraft.DesignPW:.2f} [W/kg]')
print('----------------------------------------------')
myaircraft.constraint.FindDesignPoint() print('----------------------------------------------') print(f'Design wing loading W/S: {myaircraft.DesignWTOoS:.1f} [N/m^2]') print(f'Design power-to-mass ratio P/W: {myaircraft.DesignPW:.2f} [W/kg]') print('----------------------------------------------')
---------------------------------------------- Design wing loading W/S: 3293.8 [N/m^2] Design power-to-mass ratio P/W: 194.91 [W/kg] ----------------------------------------------
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plt.plot(myaircraft.constraint.WTOoS,myaircraft.constraint.PWCruise, label='Cruise')
plt.plot(myaircraft.constraint.WTOoS,myaircraft.constraint.PWTakeOff, label='Take Off')
plt.plot(myaircraft.constraint.WTOoS,myaircraft.constraint.PWAEOClimb, label='Climb')
plt.plot(myaircraft.constraint.WTOoS,myaircraft.constraint.PWOEIClimb, label='Climb OEI')
plt.plot(myaircraft.constraint.WTOoS,myaircraft.constraint.PWTurn, label='Turn')
plt.plot(myaircraft.constraint.WTOoS,myaircraft.constraint.PWCeiling, label='Ceiling')
plt.plot(myaircraft.constraint.WTOoS,myaircraft.constraint.PWAcceleration, label='Acceleration')
plt.plot(myaircraft.constraint.WTOoSLanding,myaircraft. constraint.PWLanding, label='Landing')
plt.plot(myaircraft.DesignWTOoS, myaircraft.DesignPW, marker='o', markersize = 10, markerfacecolor = 'red', markeredgecolor = 'black')
# plt.plot(performance.WTOoSTorenbeek, performance.PWTorenbeek, label='Torenbeek')
plt.ylim([0, 300])
plt.xlim([0, 7000])
plt.legend()
plt.grid(visible=True)
plt.xlabel('$W_{TO}/S$')
plt.ylabel('$P/W_{TO}$')
plt.show()
plt.plot(myaircraft.constraint.WTOoS,myaircraft.constraint.PWCruise, label='Cruise') plt.plot(myaircraft.constraint.WTOoS,myaircraft.constraint.PWTakeOff, label='Take Off') plt.plot(myaircraft.constraint.WTOoS,myaircraft.constraint.PWAEOClimb, label='Climb') plt.plot(myaircraft.constraint.WTOoS,myaircraft.constraint.PWOEIClimb, label='Climb OEI') plt.plot(myaircraft.constraint.WTOoS,myaircraft.constraint.PWTurn, label='Turn') plt.plot(myaircraft.constraint.WTOoS,myaircraft.constraint.PWCeiling, label='Ceiling') plt.plot(myaircraft.constraint.WTOoS,myaircraft.constraint.PWAcceleration, label='Acceleration') plt.plot(myaircraft.constraint.WTOoSLanding,myaircraft. constraint.PWLanding, label='Landing') plt.plot(myaircraft.DesignWTOoS, myaircraft.DesignPW, marker='o', markersize = 10, markerfacecolor = 'red', markeredgecolor = 'black') # plt.plot(performance.WTOoSTorenbeek, performance.PWTorenbeek, label='Torenbeek') plt.ylim([0, 300]) plt.xlim([0, 7000]) plt.legend() plt.grid(visible=True) plt.xlabel('$W_{TO}/S$') plt.ylabel('$P/W_{TO}$') plt.show()
... or manually assign a wing loading
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# myaircraft.DesignWTOoS = 3323
# myaircraft.DesignWTOoS = 3323
Run Mission analysis to define the aircraft's masses
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myaircraft.weight.WeightEstimation()
myaircraft.weight.WeightEstimation()
max=58 and min=56 Optimal n 57 max=341 and min=340 Optimal n 341 max=119 and min=118 Optimal n 119 max=125 and min=124 Optimal n 125 max=125 and min=124 Optimal n 125 max=125 and min=124 Optimal n 125 max=125 and min=124 Optimal n 125
Wing surface can be readily estimated as S = WTO/(WTO/S)
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myaircraft.WingSurface = myaircraft.weight.WTO / myaircraft.DesignWTOoS * 9.81
myaircraft.WingSurface = myaircraft.weight.WTO / myaircraft.DesignWTOoS * 9.81
Well to tank impact is evaluated as a post-process
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if (myaircraft.Configuration == 'Hybrid' and WellToTankInput is not None):
myaircraft.welltowake.EvaluateSource()
if (myaircraft.Configuration == 'Hybrid' and WellToTankInput is not None): myaircraft.welltowake.EvaluateSource()
A summary of the preliminary design:
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myaircraft.Print_Aircraft_Design_Summary()
myaircraft.Print_Aircraft_Design_Summary()
Fuel mass (trip + altn + loiter): 1572.8 [Kg] Block Fuel mass: 1702.8 [Kg] Battery mass: 1686.2 [Kg] Structure: 12028.6 [Kg] Powertrain mass: 1432.9 [Kg] Empty Weight: 15647.8 [Kg] Zero Fuel Weight: 20207.8 [Kg] ---------------------------------------- Takeoff Weight: 21910.5 [Kg] Source Energy: 281499.3 [MJ] Psi: 0.0278 [-] Wing Surface: 65.3 [m^2] TakeOff engine shaft peak power: 4942.3 [KW] CLB/CRZ engine shaft peak power: 4072.2 [KW] @ 1503.5 [m] Sizing phase for thermal powertrain: Takeoff peak power Thermal powertrain rating shaft power SLS rating: 4942.3 [kW] -------------Battery Specs------------- Battery Pack Energy: 2529.4 [kWh] Battery Pack Max Power: 13563.4 [kW] Battery Pack Specific Energy: 1500.0 [Wh/kg] Battery Pack Specific Power: 8.0 [kW/kg] Battery Configuration: S-190 P-125
Tank to wake climate impact is evaluated as a post-process
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myaircraft.MissionType = 'Continue'
myaircraft.climateimpact.calculate_mission_emissions()
print(myaircraft.climateimpact.mission_emissions)
myaircraft.climateimpact.ATR()
myaircraft.MissionType = 'Continue' myaircraft.climateimpact.calculate_mission_emissions() print(myaircraft.climateimpact.mission_emissions) myaircraft.climateimpact.ATR()
{'co2': 6271.1552711333725, 'h2o': 1981.711995716157, 'so4': 0.3145574596374852, 'soot': 0.06291149192749705, 'nox': 19.691116859261136}
Out[21]:
0.0028135789146068652
Some plots:
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times = np.linspace(0,mission.profile.MissionTime2,num = 1000)
plt.plot(times/60,[mission.profile.SuppliedPowerRatio(t) for t in times], 'b')
#plt.plot(myaircraft.mission.profile.Breaks,np.ones(6)*0.05, '*')
plt.grid(visible=True)
plt.xlabel('t [min]')
plt.ylabel('Phi')
plt.show()
times = np.linspace(0,mission.profile.MissionTime2,num = 1000) plt.plot(times/60,[mission.profile.SuppliedPowerRatio(t) for t in times], 'b') #plt.plot(myaircraft.mission.profile.Breaks,np.ones(6)*0.05, '*') plt.grid(visible=True) plt.xlabel('t [min]') plt.ylabel('Phi') plt.show()
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times = np.array([])
Ef = np.array([])
Ebat = np.array([])
beta = np.array([])
soc = np.array([])
for array in mission.integral_solution:
times = np.concatenate([times, array.t])
Ef = np.concatenate([Ef , array.y[0]])
Ebat = np.concatenate([Ebat , array.y[1]])
beta = np.concatenate([beta , array.y[2]])
times = np.array([]) Ef = np.array([]) Ebat = np.array([]) beta = np.array([]) soc = np.array([]) for array in mission.integral_solution: times = np.concatenate([times, array.t]) Ef = np.concatenate([Ef , array.y[0]]) Ebat = np.concatenate([Ebat , array.y[1]]) beta = np.concatenate([beta , array.y[2]])
Takeoff propulsive power
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myaircraft.performance.TakeOff(myaircraft.DesignWTOoS,myaircraft.constraint.TakeOffConstraints['Beta'], myaircraft.constraint.TakeOffConstraints['Altitude'],myaircraft.constraint.TakeOffConstraints['kTO'], myaircraft.constraint.TakeOffConstraints['sTO'], myaircraft.constraint.DISA, myaircraft.constraint.TakeOffConstraints['Speed'], myaircraft.constraint.TakeOffConstraints['Speed Type'])
myaircraft.performance.TakeOff(myaircraft.DesignWTOoS,myaircraft.constraint.TakeOffConstraints['Beta'], myaircraft.constraint.TakeOffConstraints['Altitude'],myaircraft.constraint.TakeOffConstraints['kTO'], myaircraft.constraint.TakeOffConstraints['sTO'], myaircraft.constraint.DISA, myaircraft.constraint.TakeOffConstraints['Speed'], myaircraft.constraint.TakeOffConstraints['Speed Type'])
Out[24]:
194.88837279163596
Maximum mission Propusive Power over Takeoff Mass
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max([myaircraft.performance.PoWTO(myaircraft.DesignWTOoS,beta[t],myaircraft.mission.profile.PowerExcess(times[t]),1,myaircraft.mission.profile.Altitude(times[t]),myaircraft.mission.DISA,myaircraft.mission.profile.Velocity(times[t]),'TAS') for t in range(len(times))])
max([myaircraft.performance.PoWTO(myaircraft.DesignWTOoS,beta[t],myaircraft.mission.profile.PowerExcess(times[t]),1,myaircraft.mission.profile.Altitude(times[t]),myaircraft.mission.DISA,myaircraft.mission.profile.Velocity(times[t]),'TAS') for t in range(len(times))])
Out[25]:
160.5799464672633
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plt.plot(times/60,[myaircraft.performance.PoWTO(myaircraft.DesignWTOoS,beta[t],myaircraft.mission.profile.PowerExcess(times[t]),1,myaircraft.mission.profile.Altitude(times[t]),myaircraft.mission.DISA,myaircraft.mission.profile.Velocity(times[t]),'TAS') for t in range(len(times))], 'b')
plt.grid(visible=True)
plt.xlabel('t [min]')
plt.ylabel('Pp/W')
plt.show()
plt.plot(times/60,[myaircraft.performance.PoWTO(myaircraft.DesignWTOoS,beta[t],myaircraft.mission.profile.PowerExcess(times[t]),1,myaircraft.mission.profile.Altitude(times[t]),myaircraft.mission.DISA,myaircraft.mission.profile.Velocity(times[t]),'TAS') for t in range(len(times))], 'b') plt.grid(visible=True) plt.xlabel('t [min]') plt.ylabel('Pp/W') plt.show()
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plt.plot(times/60,[(myaircraft.weight.WTO/1000) * myaircraft.performance.PoWTO(myaircraft.DesignWTOoS,beta[t],myaircraft.mission.profile.PowerExcess(times[t]),1,myaircraft.mission.profile.Altitude(times[t]),myaircraft.mission.DISA,myaircraft.mission.profile.Velocity(times[t]),'TAS') for t in range(len(times))], 'b')
plt.grid(visible=True)
plt.xlabel('t [min]')
plt.ylabel('Pp [kW]')
plt.show()
plt.plot(times/60,[(myaircraft.weight.WTO/1000) * myaircraft.performance.PoWTO(myaircraft.DesignWTOoS,beta[t],myaircraft.mission.profile.PowerExcess(times[t]),1,myaircraft.mission.profile.Altitude(times[t]),myaircraft.mission.DISA,myaircraft.mission.profile.Velocity(times[t]),'TAS') for t in range(len(times))], 'b') plt.grid(visible=True) plt.xlabel('t [min]') plt.ylabel('Pp [kW]') plt.show()
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plt.plot(times,Ef, 'b')
#plt.plot(myaircraft.mission.profile.Breaks,np.ones(6)*0.05, '*')
plt.grid(visible=True)
plt.xlabel('t [min]')
plt.ylabel('Efuel')
plt.show()
plt.plot(times,Ef, 'b') #plt.plot(myaircraft.mission.profile.Breaks,np.ones(6)*0.05, '*') plt.grid(visible=True) plt.xlabel('t [min]') plt.ylabel('Efuel') plt.show()
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plt.plot(times,beta, 'b')
#plt.plot(myaircraft.mission.profile.Breaks,np.ones(6)*0.05, '*')
plt.grid(visible=True)
plt.xlabel('t [min]')
plt.ylabel('beta')
plt.show()
plt.plot(times,beta, 'b') #plt.plot(myaircraft.mission.profile.Breaks,np.ones(6)*0.05, '*') plt.grid(visible=True) plt.xlabel('t [min]') plt.ylabel('beta') plt.show()
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plt.plot(times/60,mission.profile.Altitude(times))
plt.grid(visible=True)
plt.xlabel('t [min]')
plt.ylabel('Altitude [m]')
plt.show()
plt.plot(times/60,mission.profile.Altitude(times)) plt.grid(visible=True) plt.xlabel('t [min]') plt.ylabel('Altitude [m]') plt.show()
