def test_mass_gain(battery, power):
print(battery)
mass_gain = find_total_mass_gain(battery)
print('mass_gain=', mass_gain)
mdot = find_mass_gain_rate(battery, power)
print('mass_gain_rate=', mdot)
return
def test_mass_gain(battery, power):
print battery
mass_gain = find_total_mass_gain(battery)
print 'mass_gain=', mass_gain
mdot = find_mass_gain_rate(battery, power)
print 'mass_gain_rate=', mdot
return
def test_mass_gain(battery,power):
print battery
mass_gain =find_total_mass_gain(battery)
print 'mass_gain=', mass_gain
mdot =find_mass_gain_rate(battery,power)
print 'mass_gain_rate=', mdot
return
def test_mass_gain(battery,power):
print(battery)
mass_gain =find_total_mass_gain(battery)
print('mass_gain=', mass_gain)
mdot =find_mass_gain_rate(battery,power)
print('mass_gain_rate=', mdot)
return
def evaluate_thrust(self,state):
# unpack
propulsor = self.propulsor
battery = self.battery
conditions = state.conditions
numerics = state.numerics
results=propulsor.evaluate_thrust(state)
Pe =results.thrust_force_vector[:,0]*conditions.freestream.velocity
try:
initial_energy=conditions.propulsion.battery_energy
if initial_energy[0][0]==0: #beginning of segment; initialize battery
battery.current_energy=battery.current_energy[-1]*np.ones_like(initial_energy)
except AttributeError: #battery energy not initialized, e.g. in takeoff
battery.current_energy=battery.current_energy[-1]*np.ones_like(F)
pbat=-Pe/self.motor_efficiency
battery_logic = Data()
battery_logic.power_in = pbat
battery_logic.current = 90. #use 90 amps as a default for now; will change this for higher fidelity methods
battery.inputs =battery_logic
battery.inputs.power_in=pbat
tol = 1e-6
battery.energy_calc(numerics)
#allow for mass gaining batteries
try:
mdot=find_mass_gain_rate(battery,-(pbat-battery.resistive_losses))
except AttributeError:
mdot=np.zeros_like(F)
#Pack the conditions for outputs
battery_draw = battery.inputs.power_in
battery_energy = battery.current_energy
conditions.propulsion.battery_draw = battery_draw
conditions.propulsion.battery_energy = battery_energy
output_power= battery_draw
#number_of_engines
#Create the outputs
results.vehicle_mass_rate = mdot
return results
def evaluate_thrust(self, state):
# unpack
propulsor = self.propulsor
battery = self.battery
conditions = state.conditions
numerics = state.numerics
results = propulsor.evaluate_thrust(state)
Pe = results.thrust_force_vector[:, 0] * conditions.freestream.velocity
try:
initial_energy = conditions.propulsion.battery_energy
if initial_energy[0][
0] == 0: #beginning of segment; initialize battery
battery.current_energy = battery.current_energy[
-1] * np.ones_like(initial_energy)
except AttributeError: #battery energy not initialized, e.g. in takeoff
battery.current_energy = battery.current_energy[-1] * np.ones_like(
F)
pbat = -Pe / self.motor_efficiency
battery_logic = Data()
battery_logic.power_in = pbat
battery_logic.current = 90. #use 90 amps as a default for now; will change this for higher fidelity methods
battery.inputs = battery_logic
battery.inputs.power_in = pbat
tol = 1e-6
battery.energy_calc(numerics)
#allow for mass gaining batteries
try:
mdot = find_mass_gain_rate(battery,
-(pbat - battery.resistive_losses))
except AttributeError:
mdot = np.zeros_like(F)
#Pack the conditions for outputs
battery_draw = battery.inputs.power_in
battery_energy = battery.current_energy
conditions.propulsion.battery_draw = battery_draw
conditions.propulsion.battery_energy = battery_energy
output_power = battery_draw
#number_of_engines
#Create the outputs
results.vehicle_mass_rate = mdot
return results
def evaluate_thrust(self, state):
""" Calculate thrust given the current state of the vehicle
Assumptions:
None
Source:
N/A
Inputs:
state [state()]
Outputs:
results.thrust_force_vector [newtons]
results.vehicle_mass_rate [kg/s]
Properties Used:
Defaulted values
"""
# unpack
propulsor = self.propulsor
battery = self.battery
conditions = state.conditions
numerics = state.numerics
results = propulsor.evaluate_thrust(state)
Pe = np.multiply(results.thrust_force_vector[:, 0],
conditions.freestream.velocity[0])
try:
initial_energy = conditions.propulsion.battery_energy
if initial_energy[0][
0] == 0: #beginning of segment; initialize battery
battery.current_energy = battery.current_energy[
-1] * np.ones_like(initial_energy)
except AttributeError: #battery energy not initialized, e.g. in takeoff
battery.current_energy = np.transpose(
np.array([battery.current_energy[-1] * np.ones_like(Pe)]))
pbat = -Pe / self.motor_efficiency
battery_logic = Data()
battery_logic.power_in = pbat
battery_logic.current = 90. #use 90 amps as a default for now; will change this for higher fidelity methods
battery.inputs = battery_logic
tol = 1e-6
battery.energy_calc(numerics)
#allow for mass gaining batteries
try:
mdot = find_mass_gain_rate(
battery,
-(pbat -
battery.resistive_losses)) #put in transpose for solver
except AttributeError:
mdot = np.zeros_like(results.thrust_force_vector[:, 0])
mdot = np.reshape(mdot, np.shape(conditions.freestream.velocity))
#Pack the conditions for outputs
battery_draw = battery.inputs.power_in
battery_energy = battery.current_energy
conditions.propulsion.battery_draw = battery_draw
conditions.propulsion.battery_energy = battery_energy
results.vehicle_mass_rate = mdot
return results
def evaluate_thrust(self,state):
""" Calculate thrust given the current state of the vehicle
Assumptions:
None
Source:
N/A
Inputs:
state [state()]
Outputs:
results.thrust_force_vector [newtons]
results.vehicle_mass_rate [kg/s]
Properties Used:
Defaulted values
"""
# unpack
propulsor = self.propulsor
battery = self.battery
conditions = state.conditions
numerics = state.numerics
results = propulsor.evaluate_thrust(state)
Pe = np.multiply(results.thrust_force_vector[:,0],conditions.freestream.velocity[0])
try:
initial_energy = conditions.propulsion.battery_energy
if initial_energy[0][0]==0: #beginning of segment; initialize battery
battery.current_energy = battery.current_energy[-1]*np.ones_like(initial_energy)
except AttributeError: #battery energy not initialized, e.g. in takeoff
battery.current_energy=np.transpose(np.array([battery.current_energy[-1]*np.ones_like(Pe)]))
pbat = -Pe/self.motor_efficiency
battery_logic = Data()
battery_logic.power_in = pbat
battery_logic.current = 90. #use 90 amps as a default for now; will change this for higher fidelity methods
battery.inputs = battery_logic
tol = 1e-6
battery.energy_calc(numerics)
#allow for mass gaining batteries
try:
mdot=find_mass_gain_rate(battery,-(pbat-battery.resistive_losses)) #put in transpose for solver
except AttributeError:
mdot=np.zeros_like(results.thrust_force_vector[:,0])
mdot=np.reshape(mdot, np.shape(conditions.freestream.velocity))
#Pack the conditions for outputs
battery_draw = battery.inputs.power_in
battery_energy = battery.current_energy
conditions.propulsion.battery_draw = battery_draw
conditions.propulsion.battery_energy = battery_energy
results.vehicle_mass_rate = mdot
return results
def evaluate_thrust(self,state):
# unpack
propulsor = self.propulsor
primary_battery = self.primary_battery
auxiliary_battery = self.auxiliary_battery
conditions = state.conditions
numerics = state.numerics
results=propulsor.evaluate_thrust(state)
Pe =results.power
try:
initial_energy = conditions.propulsion.primary_battery_energy
initial_energy_auxiliary = conditions.propulsion.auxiliary_battery_energy
if initial_energy[0][0]==0: #beginning of segment; initialize battery
primary_battery.current_energy = primary_battery.current_energy[-1]*np.ones_like(initial_energy)
auxiliary_battery.current_energy = auxiliary_battery.current_energy[-1]*np.ones_like(initial_energy)
except AttributeError: #battery energy not initialized, e.g. in takeoff
primary_battery.current_energy = np.transpose(np.array([primary_battery.current_energy[-1]*np.ones_like(Pe)]))
auxiliary_battery.current_energy = np.transpose(np.array([auxiliary_battery.current_energy[-1]*np.ones_like(Pe)]))
pbat=-Pe/self.motor_efficiency
pbat_primary=copy.copy(pbat) #prevent deep copy nonsense
pbat_auxiliary=np.zeros_like(pbat)
#print 'pbat=', pbat/10**6.
#print 'max_power prim=', primary_battery.max_power/10**6.
#print 'max_power aux=', auxiliary_battery.max_power/10**6.
for i in range(len(pbat)):
if pbat[i]<-primary_battery.max_power: #limit power output of primary_battery
pbat_primary[i] = -primary_battery.max_power #-power means discharge
pbat_auxiliary[i] = pbat[i]-pbat_primary[i]
elif pbat[i]>primary_battery.max_power: #limit charging rate of battery
pbat_primary[i] = primary_battery.max_power
pbat_auxiliary[i] = pbat[i]-pbat_primary[i]
if pbat_primary[i]>0: #don't allow non-rechargable battery to charge
pbat_primary[i] = 0
pbat_auxiliary[i] = pbat[i]
primary_battery_logic = Data()
primary_battery_logic.power_in = pbat_primary
primary_battery_logic.current = 90. #use 90 amps as a default for now; will change this for higher fidelity methods
auxiliary_battery_logic = copy.copy(primary_battery_logic)
auxiliary_battery_logic.power_in = pbat_auxiliary
primary_battery.inputs = primary_battery_logic
auxiliary_battery.inputs = auxiliary_battery_logic
tol = 1e-6
primary_battery.energy_calc(numerics)
auxiliary_battery.energy_calc(numerics)
#allow for mass gaining batteries
try:
mdot_primary = find_mass_gain_rate(primary_battery,-(pbat_primary-primary_battery.resistive_losses))
except AttributeError:
mdot_primary = np.zeros_like(results.thrust_force_vector[:,0])
try:
mdot_auxiliary = find_mass_gain_rate(auxiliary_battery,-(pbat_auxiliary-auxiliary_battery.resistive_losses))
except AttributeError:
mdot_auxiliary = np.zeros_like(results.thrust_force_vector[:,0])
mdot=mdot_primary+mdot_auxiliary
mdot=np.reshape(mdot, np.shape(conditions.freestream.velocity))
#Pack the conditions for outputs
primary_battery_draw = primary_battery.inputs.power_in
primary_battery_energy = primary_battery.current_energy
auxiliary_battery_draw = auxiliary_battery.inputs.power_in
auxiliary_battery_energy = auxiliary_battery.current_energy
conditions.propulsion.primary_battery_draw = primary_battery_draw
conditions.propulsion.primary_battery_energy = primary_battery_energy
conditions.propulsion.auxiliary_battery_draw = auxiliary_battery_draw
conditions.propulsion.auxiliary_battery_energy = auxiliary_battery_energy
results.vehicle_mass_rate = mdot
#from SUAVE.Core import Units
#print 'primary power=',primary_battery_draw/Units.MJ
#print 'aux power=',auxiliary_battery_draw/Units.MJ
return results
def evaluate_thrust(self,state):
""" Calculate thrust given the current state of the vehicle
Assumptions:
None
Source:
N/A
Inputs:
state [state()]
Outputs:
results.thrust_force_vector [newtons]
results.vehicle_mass_rate [kg/s]
conditions.propulsion.primary_battery_draw [watts]
conditions.propulsion.primary_battery_energy [joules]
conditions.propulsion.auxiliary_battery_draw [watts]
conditions.propulsion.auxiliary_battery_energy [joules]
Properties Used:
Defaulted values
"""
# unpack
propulsor = self.propulsor
primary_battery = self.primary_battery
auxiliary_battery = self.auxiliary_battery
conditions = state.conditions
numerics = state.numerics
results=propulsor.evaluate_thrust(state)
Pe=results.power
try:
initial_energy = conditions.propulsion.primary_battery_energy
initial_energy_auxiliary = conditions.propulsion.auxiliary_battery_energy
if initial_energy[0][0]==0: #beginning of segment; initialize battery
primary_battery.current_energy = primary_battery.current_energy[-1]*np.ones_like(initial_energy)
auxiliary_battery.current_energy = auxiliary_battery.current_energy[-1]*np.ones_like(initial_energy)
except AttributeError: #battery energy not initialized, e.g. in takeoff
primary_battery.current_energy = np.transpose(np.array([primary_battery.current_energy[-1]*np.ones_like(Pe)]))
auxiliary_battery.current_energy = np.transpose(np.array([auxiliary_battery.current_energy[-1]*np.ones_like(Pe)]))
pbat=-Pe/self.motor_efficiency
pbat_primary=copy.copy(pbat) #prevent deep copy nonsense
pbat_auxiliary=np.zeros_like(pbat)
for i in range(len(pbat)):
if pbat[i]<-primary_battery.max_power: #limit power output of primary_battery
pbat_primary[i] = -primary_battery.max_power #-power means discharge
pbat_auxiliary[i] = pbat[i]-pbat_primary[i]
elif pbat[i]>primary_battery.max_power: #limit charging rate of battery
pbat_primary[i] = primary_battery.max_power
pbat_auxiliary[i] = pbat[i]-pbat_primary[i]
if pbat_primary[i]>0: #don't allow non-rechargable battery to charge
pbat_primary[i] = 0
pbat_auxiliary[i] = pbat[i]
primary_battery_logic = Data()
primary_battery_logic.power_in = pbat_primary
primary_battery_logic.current = 90. #use 90 amps as a default for now; will change this for higher fidelity methods
auxiliary_battery_logic = copy.copy(primary_battery_logic)
auxiliary_battery_logic.power_in = pbat_auxiliary
primary_battery.inputs = primary_battery_logic
auxiliary_battery.inputs = auxiliary_battery_logic
tol = 1e-6
primary_battery.energy_calc(numerics)
auxiliary_battery.energy_calc(numerics)
#allow for mass gaining batteries
try:
mdot_primary = find_mass_gain_rate(primary_battery,-(pbat_primary-primary_battery.resistive_losses))
except AttributeError:
mdot_primary = np.zeros_like(results.thrust_force_vector[:,0])
try:
mdot_auxiliary = find_mass_gain_rate(auxiliary_battery,-(pbat_auxiliary-auxiliary_battery.resistive_losses))
except AttributeError:
mdot_auxiliary = np.zeros_like(results.thrust_force_vector[:,0])
mdot=mdot_primary+mdot_auxiliary
mdot=np.reshape(mdot, np.shape(conditions.freestream.velocity))
#Pack the conditions for outputs
primary_battery_draw = primary_battery.inputs.power_in
primary_battery_energy = primary_battery.current_energy
auxiliary_battery_draw = auxiliary_battery.inputs.power_in
auxiliary_battery_energy = auxiliary_battery.current_energy
conditions.propulsion.primary_battery_draw = primary_battery_draw
conditions.propulsion.primary_battery_energy = primary_battery_energy
conditions.propulsion.auxiliary_battery_draw = auxiliary_battery_draw
conditions.propulsion.auxiliary_battery_energy = auxiliary_battery_energy
results.vehicle_mass_rate = mdot
return results
def simple_sizing(interface, Ereq, Preq):
from SUAVE.Methods.Geometry.Two_Dimensional.Planform import wing_planform
# unpack
configs = interface.configs
analyses = interface.analyses
airport=interface.analyses.missions.base.airport
atmo = airport.atmosphere
mission=interface.analyses.missions.base.segments
base = configs.base
base.pull_base()
base = configs.base
base.pull_base()
Vcruise=mission['cruise'].air_speed
design_thrust=Preq*1.3/Vcruise; #approximate sealevel thrust of ducted fan
#determine geometry of fuselage as well as wings
fuselage=base.fuselages['fuselage']
SUAVE.Methods.Geometry.Two_Dimensional.Planform.fuselage_planform(fuselage)
fuselage.areas.side_projected = fuselage.heights.maximum*fuselage.lengths.cabin*1.1 # Not correct
c_vt =.08
c_ht =.6
w2v =8.54
w2h =8.54
base.wings['main_wing'] = wing_planform(base.wings['main_wing'])
SUAVE.Methods.Geometry.Two_Dimensional.Planform.vertical_tail_planform_raymer(base.wings['vertical_stabilizer'],base.wings['main_wing'], w2v, c_vt)
SUAVE.Methods.Geometry.Two_Dimensional.Planform.horizontal_tail_planform_raymer(base.wings['horizontal_stabilizer'],base.wings['main_wing'], w2h, c_ht)
base.wings['horizontal_stabilizer'] = wing_planform(base.wings['horizontal_stabilizer'])
base.wings['vertical_stabilizer'] = wing_planform(base.wings['vertical_stabilizer'])
# wing areas
for wing in base.wings:
wing.areas.wetted = 2.00 * wing.areas.reference
wing.areas.affected = 0.60 * wing.areas.reference
wing.areas.exposed = 0.75 * wing.areas.wetted
battery=base.energy_network['battery']
ducted_fan=base.propulsors['ducted_fan']
#SUAVE.Methods.Power.Battery.Sizing.initialize_from_energy_and_power(battery,Ereq,Preq)
battery.mass_properties.mass = Ereq/battery.specific_energy
battery.max_energy=Ereq
battery.max_power =Preq
battery.current_energy=[battery.max_energy] #initialize list of current energy
from SUAVE.Methods.Power.Battery.Variable_Mass import find_mass_gain_rate
m_air =-find_mass_gain_rate(battery,Ereq) #normally uses power as input to find mdot, but can use E to find m
m_water =battery.find_water_mass(Ereq)
#now add the electric motor weight
motor_mass=ducted_fan.number_of_engines*SUAVE.Methods.Weights.Correlations.Propulsion.air_cooled_motor((Preq)*Units.watts/ducted_fan.number_of_engines)
propulsion_mass=SUAVE.Methods.Weights.Correlations.Propulsion.integrated_propulsion(motor_mass/ducted_fan.number_of_engines,ducted_fan.number_of_engines)
#more geometry sizing of ducted fan
cruise_altitude= mission['climb_3'].altitude_end
conditions = atmo.compute_values(cruise_altitude)
sizing_segment = SUAVE.Components.Propulsors.Segments.Segment()
sizing_segment.M = mission['cruise'].air_speed/conditions.speed_of_sound
sizing_segment.alt = cruise_altitude
sizing_segment.T = conditions.temperature
sizing_segment.p = conditions.pressure
ducted_fan.design_thrust =design_thrust
ducted_fan.mass_properties.mass=propulsion_mass
ducted_fan.engine_sizing_ducted_fan(sizing_segment)
#compute overall mass properties
breakdown = analyses.configs.base.weights.evaluate()
breakdown.battery=battery.mass_properties.mass
breakdown.water =m_water
breakdown.air =m_air
base.mass_properties.breakdown=breakdown
#print breakdown
m_fuel=0.
base.mass_properties.operating_empty = breakdown.empty
#weight =SUAVE.Methods.Weights.Correlations.Tube_Wing.empty_custom_eng(vehicle, ducted_fan)
m_full=breakdown.empty+battery.mass_properties.mass+breakdown.payload+m_water
m_end=m_full+m_air
base.mass_properties.takeoff = m_full
base.store_diff()
# Update all configs with new base data
for config in configs:
config.pull_base()
##############################################################################
# ------------------------------------------------------------------
# Define Landing Configurations
# ------------------------------------------------------------------
landing_config=configs.landing
landing_config.wings['main_wing'].flaps.angle = 50. * Units.deg
landing_config.wings['main_wing'].slats.angle = 25. * Units.deg
landing_config.mass_properties.landing = m_end
landing_config.store_diff()
#analyses.weights=configs.base.mass_properties.takeoff
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ---------------------------------
return
