def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
# Sizing constraints for the vertical tail.
# Limiting cases: rotating torque objective (cn_beta_goal) during cruise, and
# compensation of engine failure induced torque at approach speed/altitude.
# Returns maximum area.
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs), inputs["data:geometry:propulsion:count"]
)
engine_number = inputs["data:geometry:propulsion:count"]
wing_area = inputs["data:geometry:wing:area"]
span = inputs["data:geometry:wing:span"]
l0_wing = inputs["data:geometry:wing:MAC:length"]
cg_mac_position = inputs["data:weight:aircraft:CG:aft:MAC_position"]
cn_beta_fuselage = inputs["data:aerodynamics:fuselage:cruise:CnBeta"]
cl_alpha_vt = inputs["data:aerodynamics:vertical_tail:cruise:CL_alpha"]
cruise_speed = inputs["data:TLAR:v_cruise"]
approach_speed = inputs["data:TLAR:v_approach"]
cruise_altitude = inputs["data:mission:sizing:main_route:cruise:altitude"]
wing_htp_distance = inputs["data:geometry:vertical_tail:MAC:at25percent:x:from_wingMAC25"]
nac_wet_area = inputs["data:geometry:propulsion:nacelle:wet_area"]
y_nacelle = inputs["data:geometry:propulsion:nacelle:y"]
# CASE1: OBJECTIVE TORQUE @ CRUISE #############################################################################
atm = Atmosphere(cruise_altitude)
speed_of_sound = atm.speed_of_sound
cruise_mach = cruise_speed / speed_of_sound
# Matches suggested goal by Raymer, Fig 16.20
cn_beta_goal = 0.0569 - 0.01694 * cruise_mach + 0.15904 * cruise_mach ** 2
required_cnbeta_vtp = cn_beta_goal - cn_beta_fuselage
distance_to_cg = wing_htp_distance + 0.25 * l0_wing - cg_mac_position * l0_wing
area_1 = required_cnbeta_vtp / (distance_to_cg / wing_area / span * cl_alpha_vt)
# CASE2: ENGINE FAILURE COMPENSATION DURING CLIMB ##############################################################
failure_altitude = 5000.0 # CS23 for Twin engine - at 5000ft
atm = Atmosphere(failure_altitude)
speed_of_sound = atm.speed_of_sound
pressure = atm.pressure
if engine_number == 2.0:
stall_speed = approach_speed / 1.3
mc_speed = 1.2 * stall_speed # Flights mechanics from GA - Serge Bonnet CS23
mc_mach = mc_speed / speed_of_sound
# Calculation of engine power for given conditions
flight_point = FlightPoint(
mach=mc_mach, altitude=failure_altitude, engine_setting=EngineSetting.CLIMB,
thrust_rate=1.0
) # forced to maximum thrust
propulsion_model.compute_flight_points(flight_point)
thrust = float(flight_point.thrust)
# Calculation of engine thrust and nacelle drag (failed one)
nac_drag = 0.07 * nac_wet_area # FIXME: the form factor should not be fixed outside propulsion module!
# Torque compensation
area_2 = (
2 * (y_nacelle / wing_htp_distance) * (thrust + nac_drag)
/ (pressure * mc_mach ** 2 * 0.9 * 0.42 * 10)
)
else:
area_2 = 0.0
outputs["data:geometry:vertical_tail:area"] = max(area_1, area_2)
def compute(self,
inputs,
outputs,
discrete_inputs=None,
discrete_outputs=None):
# Sizing constraints for the horizontal tail (methods from Torenbeek).
# Limiting cases: Rotating power at takeoff/landing, with the most
# forward CG position. Returns maximum area.
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs),
inputs["data:geometry:propulsion:count"])
n_engines = inputs["data:geometry:propulsion:count"]
cg_range = inputs["settings:weight:aircraft:CG:range"]
takeoff_t_rate = inputs["data:mission:sizing:takeoff:thrust_rate"]
wing_area = inputs["data:geometry:wing:area"]
x_wing_aero_center = inputs["data:geometry:wing:MAC:at25percent:x"]
lp_ht = inputs[
"data:geometry:horizontal_tail:MAC:at25percent:x:from_wingMAC25"]
wing_mac = inputs["data:geometry:wing:MAC:length"]
mtow = inputs["data:weight:aircraft:MTOW"]
mlw = inputs["data:weight:aircraft:MLW"]
x_cg_aft = inputs["data:weight:aircraft:CG:aft:x"]
z_cg_aircraft = inputs["data:weight:aircraft_empty:CG:z"]
z_cg_engine = inputs["data:weight:propulsion:engine:CG:z"]
x_lg = inputs["data:weight:airframe:landing_gear:main:CG:x"]
cl0_clean = inputs["data:aerodynamics:wing:low_speed:CL0_clean"]
cl_max_clean = inputs["data:aerodynamics:wing:low_speed:CL_max_clean"]
cl_max_landing = inputs["data:aerodynamics:aircraft:landing:CL_max"]
cl_max_takeoff = inputs["data:aerodynamics:aircraft:takeoff:CL_max"]
cl_flaps_landing = inputs["data:aerodynamics:flaps:landing:CL"]
cl_flaps_takeoff = inputs["data:aerodynamics:flaps:takeoff:CL"]
tail_efficiency_factor = inputs[
"data:aerodynamics:horizontal_tail:efficiency"]
cl_htp_landing = inputs["landing:cl_htp"]
cl_htp_takeoff = inputs["takeoff:cl_htp"]
cm_landing = inputs[
"data:aerodynamics:wing:low_speed:CM0_clean"] + inputs[
"data:aerodynamics:flaps:landing:CM"]
cm_takeoff = inputs[
"data:aerodynamics:wing:low_speed:CM0_clean"] + inputs[
"data:aerodynamics:flaps:takeoff:CM"]
cl_alpha_htp_isolated = inputs["low_speed:cl_alpha_htp_isolated"]
z_eng = z_cg_aircraft - z_cg_engine
# Conditions for calculation
atm = Atmosphere(0.0)
rho = atm.density
# CASE1: TAKE-OFF ##############################################################################################
# method extracted from Torenbeek 1982 p325
# Calculation of take-off minimum speed
weight = mtow * g
vs0 = math.sqrt(weight / (0.5 * rho * wing_area * cl_max_takeoff))
vs1 = math.sqrt(weight / (0.5 * rho * wing_area * cl_max_clean))
# Rotation speed requirement from FAR 23.51 (depends on number of engines)
if n_engines == 1:
v_r = vs1 * 1.0
else:
v_r = vs1 * 1.1
# Definition of max forward gravity center position
x_cg = x_cg_aft - cg_range * wing_mac
# Definition of horizontal tail global position
x_ht = x_wing_aero_center + lp_ht
# Calculation of wheel factor
flight_point = FlightPoint(mach=v_r / atm.speed_of_sound,
altitude=0.0,
engine_setting=EngineSetting.TAKEOFF,
thrust_rate=takeoff_t_rate)
propulsion_model.compute_flight_points(flight_point)
thrust = float(flight_point.thrust)
fact_wheel = (
(x_lg - x_cg - z_eng * thrust / weight) / wing_mac * (vs0 / v_r)**2
) # FIXME: not clear if vs0 or vs1 should be used in formula
# Compute aerodynamic coefficients for takeoff @ 0° aircraft angle
cl0_takeoff = cl0_clean + cl_flaps_takeoff
# Calculation of correction coefficient n_h and n_q
n_h = (
x_ht - x_lg
) / lp_ht * tail_efficiency_factor # tail_efficiency_factor: dynamic pressure reduction at
# tail (typical value)
n_q = 1 + cl_alpha_htp_isolated / cl_htp_takeoff * _ANG_VEL * (
x_ht - x_lg) / v_r
# Calculation of volume coefficient based on Torenbeek formula
coef_vol = (cl_max_takeoff / (n_h * n_q * cl_htp_takeoff) *
(cm_takeoff / cl_max_takeoff - fact_wheel) +
cl0_takeoff / cl_htp_takeoff *
(x_lg - x_wing_aero_center) / wing_mac)
# Calculation of equivalent area
area_1 = coef_vol * wing_area * wing_mac / lp_ht
# CASE2: LANDING ###############################################################################################
# method extracted from Torenbeek 1982 p325
# Calculation of take-off minimum speed
weight = mlw * g
vs0 = math.sqrt(weight / (0.5 * rho * wing_area * cl_max_landing))
# Rotation speed requirement from FAR 23.73
v_r = vs0 * 1.3
# Calculation of wheel factor
flight_point = FlightPoint(
mach=v_r / atm.speed_of_sound,
altitude=0.0,
engine_setting=EngineSetting.IDLE,
thrust_rate=0.1
) # FIXME: fixed thrust rate (should depend on wished descent rate)
propulsion_model.compute_flight_points(flight_point)
thrust = float(flight_point.thrust)
fact_wheel = (
(x_lg - x_cg - z_eng * thrust / weight) / wing_mac * (vs0 / v_r)**2
) # FIXME: not clear if vs0 or vs1 should be used in formula
# Evaluate aircraft overall angle (aoa)
cl0_landing = cl0_clean + cl_flaps_landing
# Calculation of correction coefficient n_h and n_q
n_h = (
x_ht - x_lg
) / lp_ht * tail_efficiency_factor # tail_efficiency_factor: dynamic pressure reduction at
# tail (typical value)
n_q = 1 + cl_alpha_htp_isolated / cl_htp_landing * _ANG_VEL * (
x_ht - x_lg) / v_r
# Calculation of volume coefficient based on Torenbeek formula
coef_vol = (cl_max_landing / (n_h * n_q * cl_htp_landing) *
(cm_landing / cl_max_landing - fact_wheel) +
cl0_landing / cl_htp_landing *
(x_lg - x_wing_aero_center) / wing_mac)
# Calculation of equivalent area
area_2 = coef_vol * wing_area * wing_mac / lp_ht
if max(area_1, area_2) < 0.0:
print(
"Warning: HTP area estimated negative (in ComputeHTArea) forced to 1m²!"
)
outputs["data:geometry:horizontal_tail:area"] = 1.0
else:
outputs["data:geometry:horizontal_tail:area"] = max(area_1, area_2)
def compute(self, inputs, outputs, discrete_inputs=None, discrete_outputs=None):
cl_max_takeoff = inputs["data:aerodynamics:wing:low_speed:CL_max_clean"]
cl0_clean = inputs["data:aerodynamics:wing:low_speed:CL0_clean"]
cl_flaps_takeoff = inputs["data:aerodynamics:flaps:takeoff:CL"]
cm_takeoff = inputs["takeoff:cm_wing"]
cl_alpha_htp_isolated = inputs["data:aerodynamics:horizontal_tail:low_speed:CL_alpha_isolated"]
cl_htp = inputs["takeoff:cl_htp"]
tail_efficiency_factor = inputs["data:aerodynamics:horizontal_tail:efficiency"]
n_engines = inputs["data:geometry:propulsion:count"]
x_wing_aero_center = inputs["data:geometry:wing:MAC:at25percent:x"]
wing_area = inputs["data:geometry:wing:area"]
wing_mac = inputs["data:geometry:wing:MAC:length"]
ht_area = inputs["data:geometry:horizontal_tail:area"]
lp_ht = inputs["data:geometry:horizontal_tail:MAC:at25percent:x:from_wingMAC25"]
mtow = inputs["data:weight:aircraft:MTOW"]
x_lg = inputs["data:weight:airframe:landing_gear:main:CG:x"]
z_cg_aircraft = inputs["data:weight:aircraft_empty:CG:z"]
z_cg_engine = inputs["data:weight:propulsion:engine:CG:z"]
propulsion_model = FuelEngineSet(
self._engine_wrapper.get_model(inputs), inputs["data:geometry:propulsion:count"]
)
# Conditions for calculation
atm = Atmosphere(0.0)
rho = atm.density
sos = atm.speed_of_sound
# Calculation of take-off minimum speed
weight = mtow * g
vs1 = math.sqrt(weight / (0.5 * rho * wing_area * cl_max_takeoff))
if n_engines == 1.0:
vr = 1.10 * vs1
else:
vr = 1.0 * vs1
mach_r = vr / sos
flight_point = FlightPoint(
mach=mach_r, altitude=0.0, engine_setting=EngineSetting.TAKEOFF,
thrust_rate=1.0
)
propulsion_model.compute_flight_points(flight_point)
thrust = float(flight_point.thrust)
x_ht = x_wing_aero_center + lp_ht
# Compute aerodynamic coefficients for takeoff @ 0° aircraft angle
cl0_takeoff = cl0_clean + cl_flaps_takeoff
eta_q = 1. + cl_alpha_htp_isolated / cl_htp * _ANG_VEL * (x_ht - x_lg) / vr
eta_h = (x_ht - x_lg) / lp_ht * tail_efficiency_factor
k_cl = cl_max_takeoff / (eta_q * eta_h * cl_htp)
tail_volume_coefficient = ht_area * lp_ht / (wing_area * wing_mac)
zt = z_cg_aircraft - z_cg_engine
engine_contribution = zt * thrust / weight
x_cg = (
1. / k_cl * (
tail_volume_coefficient - cl0_takeoff / cl_htp * (x_lg / wing_mac - 0.25)
) - cm_takeoff / cl_max_takeoff
) * (vr / vs1) ** 2.0 + x_lg - engine_contribution
outputs["data:handling_qualities:to_rotation_limit:x"] = x_cg
x_cg_ratio = (x_cg - x_wing_aero_center + 0.25 * wing_mac) / wing_mac
outputs["data:handling_qualities:to_rotation_limit:MAC_position"] = x_cg_ratio
