def climb(self, tas, alt, roc):
"""Calculate thrust during the climb.
Args:
tas (float or ndarray): True airspeed (kt).
alt (float or ndarray): Altitude(ft)
roc (float or ndarray): Vertical rate (ft/min).
Returns:
float or ndarray: Total thrust (unit: N).
"""
roc = np.abs(roc)
h = alt * aero.ft
tas = np.where(tas < 10, 10, tas)
mach = aero.tas2mach(tas * aero.kts, h)
vcas = aero.tas2cas(tas * aero.kts, h)
P = aero.pressure(h)
P10 = aero.pressure(10000 * aero.ft)
Pcr = aero.pressure(self.cruise_alt * aero.ft)
# approximate thrust at top of climb (REF 2)
Fcr = self.eng_cruise_thrust * self.eng_number
vcas_ref = aero.mach2cas(self.cruise_mach, self.cruise_alt * aero.ft)
# segment 3: alt > 30000:
d = self._dfunc(mach / self.cruise_mach)
b = (mach / self.cruise_mach)**(-0.11)
ratio_seg3 = d * np.log(P / Pcr) + b
# segment 2: 10000 < alt <= 30000:
a = (vcas / vcas_ref)**(-0.1)
n = self._nfunc(roc)
ratio_seg2 = a * (P / Pcr)**(-0.355 * (vcas / vcas_ref) + n)
# segment 1: alt <= 10000:
F10 = Fcr * a * (P10 / Pcr)**(-0.355 * (vcas / vcas_ref) + n)
m = self._mfunc(vcas / vcas_ref, roc)
ratio_seg1 = m * (P / Pcr) + (F10 / Fcr - m * (P10 / Pcr))
ratio = np.where(alt > 30000, ratio_seg3,
np.where(alt > 10000, ratio_seg2, ratio_seg1))
F = ratio * Fcr
return F
def takeoff(self, tas, alt=None):
"""Calculate thrust during the takeoff.
Args:
tas (float or ndarray): True airspeed (kt).
alt (float or ndarray): Altitude of the runway (ft). Defaults to 0.
Returns:
float or ndarray: Total thrust (unit: N).
"""
mach = aero.tas2mach(tas*aero.kts, 0)
eng_bpr = self.eng_bpr
G0 = 0.0606 * self.eng_bpr + 0.6337
if alt is None:
# at sea level
ratio = 1 - 0.377 * (1+eng_bpr) / np.sqrt((1+0.82*eng_bpr)*G0) * mach \
+ (0.23 + 0.19 * np.sqrt(eng_bpr)) * mach**2
else:
# at certain altitude
P = aero.pressure(alt * aero.ft)
dP = P / aero.p0
A = -0.4327 * dP**2 + 1.3855 * dP + 0.0472
Z = 0.9106 * dP**3 - 1.7736 * dP**2 + 1.8697 * dP
X = 0.1377 * dP**3 - 0.4374 * dP**2 + 1.3003 * dP
ratio = A - 0.377 * (1+eng_bpr) / np.sqrt((1+0.82*eng_bpr)*G0) * Z * mach \
+ (0.23 + 0.19 * np.sqrt(eng_bpr)) * X * mach**2
F = ratio * self.eng_max_thrust * self.eng_number
return F