class FuelFlow(object):
"""Fuel flow model based on ICAO emmision databank."""
def __init__(self, ac, eng=None):
"""Initialize FuelFlow object.
Args:
ac (string): ICAO aircraft type (for example: A320).
eng (string): Engine type (for example: CFM56-5A3).
Leave empty to use the default engine specified
by in the aircraft database.
"""
self.aircraft = prop.aircraft(ac)
if eng is None:
eng = self.aircraft["engine"]["default"]
self.engine = prop.engine(eng)
self.thrust = Thrust(ac, eng)
self.drag = Drag(ac)
c3, c2, c1 = (
self.engine["fuel_c3"],
self.engine["fuel_c2"],
self.engine["fuel_c1"],
)
# print(c3,c2,c1)
self.fuel_flow_model = lambda x: c3 * x**3 + c2 * x**2 + c1 * x
@ndarrayconvert
def at_thrust(self, acthr, alt=0):
"""Compute the fuel flow at a given total thrust.
Args:
acthr (int or ndarray): The total net thrust of the aircraft (unit: N).
alt (int or ndarray): Aircraft altitude (unit: ft).
Returns:
float: Fuel flow (unit: kg/s).
"""
n_eng = self.aircraft["engine"]["number"]
engthr = acthr / n_eng
ratio = engthr / self.engine["max_thrust"]
ff_sl = self.fuel_flow_model(ratio)
ff_corr_alt = self.engine["fuel_ch"] * (engthr / 1000) * (alt *
aero.ft)
ff_eng = ff_sl + ff_corr_alt
fuelflow = ff_eng * n_eng
return fuelflow
@ndarrayconvert
def takeoff(self, tas, alt=None, throttle=1):
"""Compute the fuel flow at takeoff.
The net thrust is first estimated based on the maximum thrust model
and throttle setting. Then FuelFlow.at_thrust() is called to compted
the thrust.
Args:
tas (int or ndarray): Aircraft true airspeed (unit: kt).
alt (int or ndarray): Altitude of airport (unit: ft). Defaults to sea-level.
throttle (float or ndarray): The throttle setting, between 0 and 1.
Defaults to 1, which is at full thrust.
Returns:
float: Fuel flow (unit: kg/s).
"""
Tmax = self.thrust.takeoff(tas=tas, alt=alt)
fuelflow = throttle * self.at_thrust(Tmax)
return fuelflow
@ndarrayconvert
def enroute(self, mass, tas, alt, path_angle=0):
"""Compute the fuel flow during climb, cruise, or descent.
The net thrust is first estimated based on the dynamic equation.
Then FuelFlow.at_thrust() is called to compted the thrust. Assuming
no flap deflection and no landing gear extended.
Args:
mass (int or ndarray): Aircraft mass (unit: kg).
tas (int or ndarray): Aircraft true airspeed (unit: kt).
alt (int or ndarray): Aircraft altitude (unit: ft).
path_angle (float or ndarray): Flight path angle (unit: degrees).
Returns:
float: Fuel flow (unit: kg/s).
"""
D = self.drag.clean(mass=mass, tas=tas, alt=alt, path_angle=path_angle)
# Convert angles from degrees to radians.
gamma = np.radians(path_angle)
T = D + mass * aero.g0 * np.sin(gamma)
T_idle = self.thrust.descent_idle(tas=tas, alt=alt)
T = np.where(T < 0, T_idle, T)
fuelflow = self.at_thrust(T, alt)
# do not return value outside performance boundary, with a margin of 20%
T_max = self.thrust.climb(tas=0, alt=alt, roc=0)
fuelflow = np.where(T > 1.20 * T_max, np.nan, fuelflow)
return fuelflow
def plot_model(self, plot=True):
"""Plot the engine fuel model, or return the pyplot object.
Args:
plot (bool): Display the plot or return an object.
Returns:
None or pyplot object.
"""
import matplotlib.pyplot as plt
xx = np.linspace(0, 1, 50)
yy = self.fuel_flow_model(xx)
# plt.scatter(self.x, self.y, color='k')
plt.plot(xx, yy, "--", color="gray")
if plot:
plt.show()
else:
return plt
class Performance():
"""Calulate Instantaneous performance of one object"""
def __init__(self, aircraft_name, write_output: bool = False):
self.write = write_output
# General Variables
self.aircraft_data = prop.aircraft(aircraft_name)
self.dt = 1.0 / 60.0 # simulation timestep 60 per seconds
aircraft_txt = open(f"./data/{aircraft_name}.json", 'r').read()
self.aircraft = json.loads(aircraft_txt)
eng_name = self.aircraft_data["engine"]["default"]
self.ac_thrust = Thrust(ac=self.aircraft["Name"], eng=eng_name)
# self.ac_thrust = Thrust(f"./data/{aircraft_name}_thrust.csv")
# Performance Variables (all unit in SI except stated otherwise)
self.lift_drag = LiftDrag(f"./data/{aircraft_name}_ld.csv")
self.g = 9.81
self.mass = self.aircraft_data["limits"]["MTOW"]
self.thrust_lever = 1.0
self.altitude = 0.0
self.pressure = 0.0
self.density = 0.0
self.temp = 0.0
self.cas = 0.0
self.tas = 0.0
self.v_y = 0.0 # vertical Speed
self.vs = 0.0 # Vertical Speed [fpm]
self.drag = 0.0
self.thrust = 0.0
self.lift = 0.0
self.weight = 0.0
self.t_d = 0.0 # thrust minus drag aka exceed thrust
self.l_w = 0.0 # lift minus weight aka exceed lift
self.pitch = 0.0
self.fpa = 0.0
self.aoa = 0.0
self.Q = 0.0 # tas² * density
self.acc_x = 0.0
self.acc_y = 0.0
self.distance_x = 0.0 # total distance[m]
self.d_x = 0.0 # instantaneous X distance[m]
self.d_y = 0.0 # instantaneous Y distance[m]
self.phase = 0 # Current phase
self.cd = 0.0
self.cl = 0.0
self.drag0 = self.aircraft["Drag0"]
self.lift0 = self.aircraft["Lift0"]
self.gear = False
self.flaps = 0
self.pitch_target = 0.0
self.pitch_rate_of_change = 3.0 # rate of change of the pitch [°/sec]
self.ac_fuelflow = FuelFlow(ac=self.aircraft["Name"], eng=eng_name)
if self.write:
self.output = open("output.csv", 'w+')
self.output.write(self.__get_header())
self.output.flush()
def __get_Q(self):
self.pressure, self.density, self.temp = aero.atmos(self.altitude)
self.Q = self.density * (self.tas**2)
def __calculate_FPA(self):
self.fpa = math.degrees(math.atan2(self.d_y, self.d_x))
self.aoa = self.pitch - self.fpa
def __calculate_lift(self):
if self.gear:
self.cl += self.aircraft["Gear"]["Lift"]
if self.flaps > 0:
self.cl += self.aircraft["Flaps"][self.flaps - 1]["Lift"]
self.lift = self.lift0\
+ (0.5 * self.Q * self.aircraft["WingSpan"] * self.cl)
self.lift *= 1.5
def __calculate_drag(self, new: bool = True):
if self.gear:
self.cd += self.aircraft["Gear"]["Drag"]
if self.flaps > 0:
self.cd += self.aircraft["Flaps"][self.flaps - 1]["Drag"]
self.drag = self.drag0\
+ (0.5 * self.Q * self.aircraft["WingSpan"] * self.cd)
def __change_pitch(self) -> None:
"""
Change pitch in relation of pitch target change of pitch occure with
a 3 degrees per seconds change.
"""
if self.pitch == self.pitch_target:
return
if self.pitch > self.pitch_target:
self.pitch -= self.pitch_rate_of_change * self.dt
elif self.pitch < self.pitch_target:
self.pitch += self.pitch_rate_of_change * self.dt
def run(self) -> bool:
"""Calculate aircraft performance till thrust reduction altitude
Args:
target_alt (float, optional):
The thrust reduction altitude in meters.
Defaults to 457.2m (1500.0 feets)
Returns:
bool: True if the phase is still valid, else False
"""
if self.distance_x == 0:
self.aoa = self.pitch
self.cl, self.cd = self.lift_drag.get_data(self.aoa)
self.__get_Q()
self.__change_pitch()
self.__calculate_drag()
self.__calculate_lift()
self.g = local_gravity(50.0, self.altitude)
self.weight = self.mass * self.g
max_thrust = self.ac_thrust.takeoff(alt=self.altitude / aero.ft,
tas=self.tas / aero.kts)
idle_thrust = self.ac_thrust.descent_idle(tas=self.tas / aero.kts,
alt=self.altitude / aero.ft)
self.thrust = interpolate(self.thrust_lever, 0.0, 1.0, idle_thrust,
max_thrust)
fuelflow = self.ac_fuelflow.at_thrust(acthr=self.thrust / 2.0,
alt=self.altitude / aero.ft)
self.mass -= fuelflow * self.dt * 2
self.t_d = self.thrust - self.drag\
- (self.weight * math.sin(math.radians(self.pitch)))
self.l_w = self.lift\
- (self.weight * math.cos(math.radians(self.pitch)))\
+ (self.thrust * math.sin(math.radians(self.pitch)))
acc = self.t_d / self.mass
self.acc_x = acc * math.cos(math.radians(self.pitch))
self.acc_y = acc * math.sin(math.radians(self.pitch))
v_acc = self.l_w / self.mass
self.acc_y += v_acc * math.cos(math.radians(self.pitch))
self.acc_x += v_acc * math.sin(math.radians(self.pitch))
self.d_x = (self.tas * self.dt) + 0.5 * self.acc_x * (self.dt**2)
self.d_y = (self.v_y * self.dt) + 0.5 * self.acc_y * (self.dt**2)
self.tas += self.acc_x * self.dt
self.cas = aero.tas2cas(self.tas, self.altitude)
self.v_y += self.acc_y * self.dt
if self.altitude <= 0:
self.altitude = 0
if self.d_y < 0:
self.d_y = 0
if self.v_y < 0:
self.v_y = 0
self.vs = 0
self.altitude += self.d_y
self.distance_x += self.d_x
self.vs = self.v_y / aero.fpm
self.__calculate_FPA()
if self.write:
self.output.write(str(self))
self.output.flush()
def __str__(self):
return f"{self.mass},{self.altitude},{self.pressure},{self.density},"\
f"{self.temp},{self.cas},{self.tas},{self.v_y},{self.vs},"\
f"{self.drag},{self.thrust},{self.t_d},{self.pitch},"\
f"{self.fpa},{self.aoa},{self.Q},{self.acc_x},{self.acc_y},"\
f"{self.distance_x},{self.d_x},{self.d_y},{self.phase},"\
f"{self.cd},{self.drag0},{self.gear},{self.flaps},{self.cl},"\
f"{self.lift},{self.weight},{self.l_w},{self.thrust_lever},"\
f"{self.altitude / aero.ft},{self.g},{self.pitch_target}\n"
def __get_header(self):
return "Mass,Altitude,Pressure,Density,Temperature,Cas,Tas,Vy,VS,"\
"Drag,Thrust,T-D,Pitch,FPA,AOA,Q,AccelerationX,AccelerationY,"\
"DistanceX,Dx,Dy,Phase,Cd,Cd0,Gear,Flaps,Cl,Lift,Weight,L-W,"\
"Thrust Limit,Altitude FT,Gravity,Target Pitch\n"
class FuelFlow(object):
"""Fuel flow model based on ICAO emmision databank."""
def __init__(self, ac, eng=None):
"""Initialize FuelFlow object.
Args:
ac (string): ICAO aircraft type (for example: A320).
eng (string): Engine type (for example: CFM56-5A3).
Leave empty to use the default engine specified
by in the aircraft database.
"""
self.aircraft = prop.aircraft(ac)
if eng is None:
eng = self.aircraft['engine']['default']
self.engine = prop.engine(eng)
self.thrust = Thrust(ac, eng)
self.drag = Drag(ac)
c3, c2, c1 = self.engine['fuel_c3'], self.engine[
'fuel_c2'], self.engine['fuel_c1']
# print(c3,c2,c1)
self.fuel_flow_model = lambda x: c3 * x**3 + c2 * x**2 + c1 * x
@ndarrayconvert
def at_thrust(self, acthr, alt=0):
"""Compute the fuel flow at a given total thrust.
Args:
acthr (int or ndarray): The total net thrust of the aircraft (unit: N).
alt (int or ndarray): Aicraft altitude (unit: ft).
Returns:
float: Fuel flow (unit: kg/s).
"""
ratio = acthr / (self.engine['max_thrust'] *
self.aircraft['engine']['number'])
fuelflow = self.fuel_flow_model(ratio) * self.aircraft['engine']['number'] \
+ self.engine['fuel_ch'] * (alt*aero.ft) * (acthr/1000)
return fuelflow
@ndarrayconvert
def takeoff(self, tas, alt=None, throttle=1):
"""Compute the fuel flow at takeoff.
The net thrust is first estimated based on the maximum thrust model
and throttle setting. Then FuelFlow.at_thrust() is called to compted
the thrust.
Args:
tas (int or ndarray): Aircraft true airspeed (unit: kt).
alt (int or ndarray): Altitude of airport (unit: ft). Defaults to sea-level.
throttle (float or ndarray): The throttle setting, between 0 and 1.
Defaults to 1, which is at full thrust.
Returns:
float: Fuel flow (unit: kg/s).
"""
Tmax = self.thrust.takeoff(tas=tas, alt=alt)
fuelflow = throttle * self.at_thrust(Tmax)
return fuelflow
@ndarrayconvert
def enroute(self, mass, tas, alt, path_angle=0):
"""Compute the fuel flow during climb, cruise, or descent.
The net thrust is first estimated based on the dynamic equation.
Then FuelFlow.at_thrust() is called to compted the thrust. Assuming
no flap deflection and no landing gear extended.
Args:
mass (int or ndarray): Aircraft mass (unit: kg).
tas (int or ndarray): Aircraft true airspeed (unit: kt).
alt (int or ndarray): Aircraft altitude (unit: ft).
path_angle (float or ndarray): Flight path angle (unit: ft).
Returns:
float: Fuel flow (unit: kg/s).
"""
D = self.drag.clean(mass=mass, tas=tas, alt=alt, path_angle=path_angle)
gamma = np.radians(path_angle)
T = D + mass * aero.g0 * np.sin(gamma)
T_idle = self.thrust.descent_idle(tas=tas, alt=alt)
T = np.where(T < 0, T_idle, T)
fuelflow = self.at_thrust(T, alt)
return fuelflow
def plot_model(self, plot=True):
"""Plot the engine fuel model, or return the pyplot object.
Args:
plot (bool): Display the plot or return an object.
Returns:
None or pyplot object.
"""
import matplotlib.pyplot as plt
xx = np.linspace(0, 1, 50)
yy = self.fuel_flow_model(xx)
# plt.scatter(self.x, self.y, color='k')
plt.plot(xx, yy, '--', color='gray')
if plot:
plt.show()
else:
return plt