def update(self, simt=1):
super(OpenAP, self).update(simt)
# update phase, infer from spd, roc, alt
self.phase = ph.get(self.lifttype,
bs.traf.tas,
bs.traf.vs,
bs.traf.alt,
unit='SI')
# update limits, based on phase change
limits = self.__construct_limit_matrix(self.actypes, self.phase)
self.vmin = limits[:, 0]
self.vmax = limits[:, 1]
self.vsmin = limits[:, 2]
self.vsmax = limits[:, 3]
self.hmax = limits[:, 4]
self.axmax = limits[:, 5]
# compute thrust
# = number of engines x engine static thrust x thrust ratio
idx_fixwing = np.where(self.lifttype == coeff.LIFT_FIXWING)[0]
self.thrustratio[idx_fixwing] = thrust.compute_thrust_ratio(
self.phase[idx_fixwing], self.engbpr[idx_fixwing],
bs.traf.tas[idx_fixwing], bs.traf.alt[idx_fixwing])
self.thrust[idx_fixwing] = self.engnum[idx_fixwing] * self.engthrust[
idx_fixwing] * self.thrustratio[idx_fixwing]
# compute fuel flow
self.fuelflow = self.engnum * (self.ff_coeff_a * self.thrustratio**2 \
+ self.ff_coeff_b * self.thrustratio \
+ self.ff_coeff_c)
# TODO: implement thrust computation for rotor aircraft
# idx_rotor = np.where(self.lifttype==coeff.LIFT_ROTOR)[0]
# self.thrust[idx_rotor] = 0
# update bank angle, due to phase change
self.bank = np.where((self.phase == ph.TO) | (self.phase == ph.LD), 15,
self.bank)
self.bank = np.where((self.phase == ph.IC) | (self.phase == ph.CR) |
(self.phase == ph.AP), 35, self.bank)
return None
def update(self, dt=bs.settings.performance_dt):
super(OpenAP, self).update()
# update phase, infer from spd, roc, alt
lenph1 = len(self.phase)
self.phase = ph.get(self.lifttype,
bs.traf.tas,
bs.traf.vs,
bs.traf.alt,
unit="SI")
# update speed limits, based on phase change
self.vmin, self.vmax = self._construct_v_limits(
self.actypes, self.phase)
idx_fixwing = np.where(self.lifttype == coeff.LIFT_FIXWING)[0]
# ----- compute drag -----
# update drage coefficient based on flight phase
self.cd0[self.phase == ph.GD] = (
self.cd0_to[self.phase == ph.GD] +
self.delta_cd_gear[self.phase == ph.GD])
self.cd0[self.phase == ph.IC] = self.cd0_to[self.phase == ph.IC]
self.cd0[self.phase == ph.AP] = self.cd0_ld[self.phase == ph.AP]
self.cd0[self.phase == ph.CL] = self.cd0_clean[self.phase == ph.CL]
self.cd0[self.phase == ph.CR] = self.cd0_clean[self.phase == ph.CR]
self.cd0[self.phase == ph.DE] = self.cd0_clean[self.phase == ph.DE]
self.cd0[self.phase == ph.NA] = self.cd0_clean[self.phase == ph.NA]
self.k[self.phase == ph.GD] = self.k_to[self.phase == ph.GD]
self.k[self.phase == ph.IC] = self.k_to[self.phase == ph.IC]
self.k[self.phase == ph.AP] = self.k_ld[self.phase == ph.AP]
self.k[self.phase == ph.CL] = self.k_clean[self.phase == ph.CL]
self.k[self.phase == ph.CR] = self.k_clean[self.phase == ph.CR]
self.k[self.phase == ph.DE] = self.k_clean[self.phase == ph.DE]
self.k[self.phase == ph.NA] = self.k_clean[self.phase == ph.NA]
rho = aero.vdensity(bs.traf.alt[idx_fixwing])
vtas = bs.traf.tas[idx_fixwing]
rhovs = 0.5 * rho * vtas**2 * self.Sref[idx_fixwing]
cl = self.mass[idx_fixwing] * aero.g0 / rhovs
self.drag[idx_fixwing] = rhovs * (self.cd0[idx_fixwing] +
self.k[idx_fixwing] * cl**2)
# ----- compute maximum thrust -----
max_thrustratio_fixwing = thrust.compute_max_thr_ratio(
self.phase[idx_fixwing],
self.engbpr[idx_fixwing],
bs.traf.tas[idx_fixwing],
bs.traf.alt[idx_fixwing],
bs.traf.vs[idx_fixwing],
self.engnum[idx_fixwing] * self.engthrmax[idx_fixwing],
)
self.max_thrust[idx_fixwing] = (max_thrustratio_fixwing *
self.engnum[idx_fixwing] *
self.engthrmax[idx_fixwing])
# ----- compute net thrust -----
self.thrust[idx_fixwing] = (
self.drag[idx_fixwing] +
self.mass[idx_fixwing] * bs.traf.ax[idx_fixwing])
# ----- compute duel flow -----
thrustratio_fixwing = self.thrust[idx_fixwing] / (
self.engnum[idx_fixwing] * self.engthrmax[idx_fixwing])
self.fuelflow[idx_fixwing] = self.engnum[idx_fixwing] * (
self.ff_coeff_a[idx_fixwing] * thrustratio_fixwing**2 +
self.ff_coeff_b[idx_fixwing] * thrustratio_fixwing +
self.ff_coeff_c[idx_fixwing])
# TODO: implement thrust computation for rotor aircraft
# idx_rotor = np.where(self.lifttype==coeff.LIFT_ROTOR)[0]
# self.thrust[idx_rotor] = 0
# update bank angle, due to phase change
self.bank = np.where((self.phase == ph.GD), 15, self.bank)
self.bank = np.where(
(self.phase == ph.IC) | (self.phase == ph.CR) |
(self.phase == ph.AP),
35,
self.bank,
)
# ----- debug statements -----
# print(bs.traf.id)
# print(self.phase)
# print(self.thrust.astype(int))
# print(np.round(self.fuelflow, 2))
# print(self.drag.astype(int))
# print()
return None
def update(self, dt=bs.settings.performance_dt):
super(OpenAP, self).update()
# update phase, infer from spd, roc, alt
lenph1 = len(self.phase)
self.phase = ph.get(self.lifttype, bs.traf.tas, bs.traf.vs, bs.traf.alt, unit='SI')
# update limits, based on phase change
limits = self.__construct_limit_matrix(self.actypes, self.phase)
self.vmin = limits[:, 0]
self.vmax = limits[:, 1]
self.vsmin = limits[:, 2]
self.vsmax = limits[:, 3]
self.hmax = limits[:, 4]
self.axmax = limits[:, 5]
self.vminto = limits[:, 6]
idx_fixwing = np.where(self.lifttype==coeff.LIFT_FIXWING)[0]
# ----- compute drag -----
# update drage coefficient based on flight phase
self.cd0[self.phase==ph.GD] = self.cd0_to[self.phase==ph.GD]
self.cd0[self.phase==ph.IC] = self.cd0_ic[self.phase==ph.IC]
self.cd0[self.phase==ph.AP] = self.cd0_ap[self.phase==ph.AP]
self.cd0[self.phase==ph.CL] = self.cd0_clean[self.phase==ph.CL]
self.cd0[self.phase==ph.CR] = self.cd0_clean[self.phase==ph.CR]
self.cd0[self.phase==ph.DE] = self.cd0_clean[self.phase==ph.DE]
self.cd0[self.phase==ph.NA] = self.cd0_clean[self.phase==ph.NA]
rho = aero.vdensity(bs.traf.alt[idx_fixwing])
vtas = bs.traf.tas[idx_fixwing]
rhovs = 0.5 * rho * vtas**2 * self.Sref[idx_fixwing]
cl = self.mass[idx_fixwing] * aero.g0 / rhovs
self.drag[idx_fixwing] = rhovs * (self.cd0[idx_fixwing] + self.k[idx_fixwing] * cl**2)
# ----- compute maximum thrust -----
max_thrustratio_fixwing = thrust.compute_max_thr_ratio(
self.phase[idx_fixwing], self.engbpr[idx_fixwing],
bs.traf.tas[idx_fixwing], bs.traf.alt[idx_fixwing],
bs.traf.vs[idx_fixwing], self.engnum[idx_fixwing]*self.engthrmax[idx_fixwing]
)
self.max_thrust[idx_fixwing] = max_thrustratio_fixwing * self.engnum[idx_fixwing] * self.engthrmax[idx_fixwing]
# ----- compute net thrust -----
self.thrust[idx_fixwing] = self.drag[idx_fixwing] + self.mass[idx_fixwing] * bs.traf.ax[idx_fixwing]
# ----- compute duel flow -----
thrustratio_fixwing = self.thrust[idx_fixwing] / (self.engnum[idx_fixwing] * self.engthrmax[idx_fixwing])
self.fuelflow[idx_fixwing] = self.engnum[idx_fixwing] * (self.ff_coeff_a[idx_fixwing] * thrustratio_fixwing**2 \
+ self.ff_coeff_b[idx_fixwing] * thrustratio_fixwing \
+ self.ff_coeff_c[idx_fixwing])
# TODO: implement thrust computation for rotor aircraft
# idx_rotor = np.where(self.lifttype==coeff.LIFT_ROTOR)[0]
# self.thrust[idx_rotor] = 0
# update bank angle, due to phase change
self.bank = np.where((self.phase==ph.GD), 15, self.bank)
self.bank = np.where((self.phase==ph.IC) | (self.phase==ph.CR) | (self.phase==ph.AP), 35, self.bank)
# ----- debug statements -----
# print(bs.traf.id)
# print(self.phase)
# print(self.thrust.astype(int))
# print(np.round(self.fuelflow, 2))
# print(self.drag.astype(int))
# print()
return None
def update(self, simt=1):
super(OpenAP, self).update(simt)
# update phase, infer from spd, roc, alt
lenph1 = len(self.phase)
self.phase = ph.get(self.lifttype,
bs.traf.tas,
bs.traf.vs,
bs.traf.alt,
unit='SI')
# update limits, based on phase change
limits = self.__construct_limit_matrix(self.actypes, self.phase)
self.vmin = limits[:, 0]
self.vmax = limits[:, 1]
self.vsmin = limits[:, 2]
self.vsmax = limits[:, 3]
self.hmax = limits[:, 4]
self.axmax = limits[:, 5]
idx_fixwing = np.where(self.lifttype == coeff.LIFT_FIXWING)[0]
# ----- compute drag -----
# update drage coefficient based on flight phase
self.cd0[self.phase == ph.TO] = self.cd0_to[self.phase == ph.TO]
self.cd0[self.phase == ph.IC] = self.cd0_ic[self.phase == ph.IC]
self.cd0[self.phase == ph.AP] = self.cd0_ap[self.phase == ph.AP]
self.cd0[self.phase == ph.LD] = self.cd0_ld[self.phase == ph.LD]
self.cd0[self.phase == ph.GD] = self.cd0_gd[self.phase == ph.GD]
self.cd0[self.phase == ph.CL] = self.cd0_clean[self.phase == ph.CL]
self.cd0[self.phase == ph.CR] = self.cd0_clean[self.phase == ph.CR]
self.cd0[self.phase == ph.DE] = self.cd0_clean[self.phase == ph.DE]
self.cd0[self.phase == ph.NA] = self.cd0_clean[self.phase == ph.NA]
rho = aero.vdensity(bs.traf.alt[idx_fixwing])
vtas = bs.traf.tas[idx_fixwing]
rhovs = 0.5 * rho * vtas**2 * self.Sref[idx_fixwing]
cl = self.mass[idx_fixwing] * aero.g0 / rhovs
self.drag[idx_fixwing] = rhovs * (self.cd0[idx_fixwing] +
self.k[idx_fixwing] * cl**2)
# ----- compute thrust -----
# eq: number of engines x engine static thrust x thrust ratio
self.thrustratio[idx_fixwing] = thrust.compute_thrust_ratio(
self.phase[idx_fixwing], self.engbpr[idx_fixwing],
bs.traf.tas[idx_fixwing], bs.traf.alt[idx_fixwing],
bs.traf.vs[idx_fixwing],
self.engnum[idx_fixwing] * self.engthrust[idx_fixwing])
self.thrust[idx_fixwing] = self.engnum[idx_fixwing] * self.engthrust[
idx_fixwing] * self.thrustratio[idx_fixwing]
# ----- compute duel flow -----
self.fuelflow = self.engnum * (self.ff_coeff_a * self.thrustratio**2 \
+ self.ff_coeff_b * self.thrustratio \
+ self.ff_coeff_c)
# TODO: implement thrust computation for rotor aircraft
# idx_rotor = np.where(self.lifttype==coeff.LIFT_ROTOR)[0]
# self.thrust[idx_rotor] = 0
# update bank angle, due to phase change
self.bank = np.where((self.phase == ph.TO) | (self.phase == ph.LD), 15,
self.bank)
self.bank = np.where((self.phase == ph.IC) | (self.phase == ph.CR) |
(self.phase == ph.AP), 35, self.bank)
# ----- debug statements -----
# print(bs.traf.id)
# print(self.phase)
# print(self.thrust.astype(int))
# print(np.round(self.fuelflow, 2))
# print(self.drag.astype(int))
# print()
return None
