def air_path(aircraft, nei, altp, disa, speed_mode, speed, mass, rating):
"""
Retrieves air path in various conditions
"""
g = earth.gravity()
[pamb, tamb, tstd, dtodz] = earth.atmosphere(altp, disa)
mach = get_mach(pamb, speed_mode, speed)
fn, data = propu.thrust(aircraft, pamb, tamb, mach, rating, nei)
cz = lift_from_speed(aircraft, pamb, mach, mass)
[cx, lod] = aero.drag(aircraft, pamb, tamb, mach, cz)
if (nei > 0):
dcx = propu.oei_drag(aircraft, pamb, mach)
cx = cx + dcx * nei
lod = cz / cx
acc_factor = earth.climb_mode(speed_mode, dtodz, tstd, disa, mach)
slope = (fn / (mass * g) - 1 / lod) / acc_factor
vsnd = earth.sound_speed(tamb)
v_z = mach * vsnd * slope
return slope, v_z
def acceleration(aircraft, nei, altp, disa, speed_mode, speed, mass, rating):
"""
Aircraft acceleration on level flight
"""
wing = aircraft.wing
gam = earth.heat_ratio()
[pamb, tamb, tstd, dtodz] = earth.atmosphere(altp, disa)
mach = get_mach(pamb, speed_mode, speed)
fn, Data = propu.thrust(aircraft, pamb, tamb, mach, rating, nei)
cz = lift_from_speed(aircraft, pamb, mach, mass)
cx, lod = aero.drag(aircraft, pamb, tamb, mach, cz)
if (nei > 0):
dcx = propu.oei_drag(aircraft, pamb, mach)
cx = cx + dcx * nei
acc = (fn - 0.5 * gam * pamb * mach**2 * wing.area * cx) / mass
return acc
def take_off(aircraft, kvs1g, altp, disa, mass, hld_conf):
"""
Take off field length and climb path at 35 ft depending on stall margin (kVs1g)
"""
wing = aircraft.wing
propulsion = aircraft.propulsion
(MTO, MCN, MCL, MCR, FID) = propulsion.rating_code
czmax, cz_0 = craft_aero.high_lift(wing, hld_conf)
rating = MTO
[pamb, tamb, tstd, dtodz] = earth.atmosphere(altp, disa)
[rho, sig] = earth.air_density(pamb, tamb)
cz_to = czmax / kvs1g**2
mach = flight.speed_from_lift(aircraft, pamb, cz_to, mass)
nei = 0 # For Magic Line factor computation
fn, trash = propu.thrust(aircraft, pamb, tamb, mach, rating, nei)
ml_factor = mass**2 / (cz_to * fn * wing.area * sig**0.8
) # Magic Line factor
tofl = 14.23 * ml_factor + 97.58
nei = 1 # For 2nd segment computation
speed_mode = 1
speed = flight.get_speed(pamb, speed_mode, mach)
seg2path, vz = flight.air_path(aircraft, nei, altp, disa, speed_mode,
speed, mass, rating)
return seg2path, tofl
#------------------------------------------------------------------------------------------------------
run.mass_mission_adaptation(aircraft)
#run.mass_estimation(aircraft)
# Calculate all airplane performances
#------------------------------------------------------------------------------------------------------
run.performance_analysis(aircraft)
# Print relevant output
#------------------------------------------------------------------------------------------------------
altp = unit.m_ft(35000)
disa = 0
pamb, tamb, tstd, dtodz = earth.atmosphere(altp, disa)
(MTO, MCN, MCL, MCR, FID) = aircraft.propulsion.rating_code
nei = 0
Fn, Data = propu.thrust(aircraft, pamb, tamb, cruise_mach, MTO, nei)
vsnd = earth.sound_speed(tamb)
tas = vsnd * cruise_mach
print("")
print("True air speed in cruise", "%.1f" % tas, " m/s")
print("Totalthrust in cruise", "%.0f" % Fn, " N")
print("")
print("Engine thrust = ",
"%.1f" % (aircraft.propulsion.reference_thrust_effective / 10), " daN")
print("Wing area = ", "%.1f" % aircraft.wing.area, " m2")
print("MTOW = ", "%.0f" % aircraft.weights.mtow, " kg")
print("OWE = ", "%.0f" % aircraft.weights.owe, " kg")
print("")
print("Turbofan nacelle width = ", "%.1f" % aircraft.turbofan_nacelle.width,
def eval_propulsion_design(aircraft):
"""
Propulsion architecture design
"""
propulsion = aircraft.propulsion
propulsion.rating_code = (0, 1, 2, 3, 4)
if (propulsion.architecture == 1):
engine = aircraft.turbofan_engine
eval_turbofan_engine_design(aircraft)
eval_turbofan_nacelle_design(aircraft)
elif (propulsion.architecture == 2):
engine = aircraft.turbofan_engine
eval_turbofan_engine_design(aircraft)
eval_hybrid_engine_design(aircraft)
eval_hybrid_nacelle_design(aircraft)
elif (propulsion.architecture == 3):
engine = aircraft.turbofan_engine
eval_hybrid_turbofan_engine_design(aircraft)
eval_hybrid_engine_design(aircraft)
eval_hybrid_body_nacelle_design(aircraft)
elif (propulsion.architecture == 4):
engine = aircraft.turboprop_engine
eval_turboprop_engine_design(aircraft)
eval_turboprop_nacelle_design(aircraft)
else:
raise Exception("propulsion.architecture index is out of range")
(MTO, MCN, MCL, MCR, FID) = propulsion.rating_code
disa = 15.
altp = 0.
mach = 0.25
nei = 0.
(pamb, tamb, tstd, dtodz) = earth.atmosphere(altp, disa)
(Fn, Data) = propu.thrust(aircraft, pamb, tamb, mach, MTO, nei)
propulsion.reference_thrust_effective = (Fn / engine.n_engine) / 0.80
propulsion.mto_thrust_ref = Fn / engine.n_engine
disa = aircraft.low_speed.disa_oei
altp = aircraft.low_speed.req_oei_altp
mach = 0.5 * aircraft.design_driver.cruise_mach
nei = 1.
(pamb, tamb, tstd, dtodz) = earth.atmosphere(altp, disa)
(Fn, Data) = propu.thrust(aircraft, pamb, tamb, mach, MCN, nei)
propulsion.mcn_thrust_ref = Fn / (engine.n_engine - nei)
disa = 0.
altp = aircraft.design_driver.ref_cruise_altp
mach = aircraft.design_driver.cruise_mach
nei = 0.
(pamb, tamb, tstd, dtodz) = earth.atmosphere(altp, disa)
propulsion.sfc_cruise_ref = propu.sfc(aircraft, pamb, tamb, mach, MCR, nei)
if (propulsion.architecture == 1):
sec = 0.
elif (propulsion.architecture == 2):
fn, sec, data = propu.hybrid_thrust(aircraft, pamb, tamb, mach, MCR,
nei)
elif (propulsion.architecture == 3):
fn, sec, data = propu.hybrid_thrust(aircraft, pamb, tamb, mach, MCR,
nei)
elif (propulsion.architecture == 4):
sec = 0.
else:
raise Exception("propulsion.architecture index is out of range")
propulsion.sec_cruise_ref = sec
(Fn, Data) = propu.thrust(aircraft, pamb, tamb, mach, FID, nei)
propulsion.fid_thrust_ref = Fn / engine.n_engine
disa = 0.
altp = aircraft.design_driver.top_of_climb_altp
mach = aircraft.design_driver.cruise_mach
nei = 0.
(pamb, tamb, tstd, dtodz) = earth.atmosphere(altp, disa)
(Fn, Data) = propu.thrust(aircraft, pamb, tamb, mach, MCL, nei)
propulsion.mcl_thrust_ref = Fn / engine.n_engine
(Fn, Data) = propu.thrust(aircraft, pamb, tamb, mach, MCR, nei)
propulsion.mcr_thrust_ref = Fn / engine.n_engine
return
def vertical_tail_sizing(aircraft):
"""
Computes necessary VTP area variation to meet engine failure case constraint
Influence of CG position is ignored
"""
design_driver = aircraft.design_driver
propulsion = aircraft.propulsion
wing = aircraft.wing
vtp = aircraft.vertical_tail
aerodynamics = aircraft.aerodynamics
payload = aircraft.payload
c_of_g = aircraft.center_of_gravity
(MTO, MCN, MCL, MCR, FID) = propulsion.rating_code
cyb_vtp, xlc_vtp, aoa_max_vtp, ki_vtp = frame_aero.vtp_aero_data(aircraft)
payload = 0.5 * payload.nominal # Light payload
range = design_driver.design_range / 15 # Short mission
altp = design_driver.ref_cruise_altp
mach = design_driver.cruise_mach
disa = 30 # Hot condition
tow, block_fuel, block_time, total_fuel = sub_proc.mission_tow(
aircraft, payload, range, altp, mach, disa)
altp = 0
disa = 15
pamb, tamb, tstd, dtodz = earth.atmosphere(altp, disa)
stall_margin = regul.kvs1g_min_take_off()
czmax_to = aerodynamics.cz_max_to
mach_s1g = flight.speed_from_lift(aircraft, pamb, czmax_to, tow)
mach_35ft = stall_margin * mach_s1g # V2 speed
mach_mca = mach_35ft / 1.1 #Approximation of required VMCA
altp = 0
disa = 15
nei = 1
pamb, tamb, tstd, dtodz = earth.atmosphere(altp, disa)
fn, data = propu.thrust(aircraft, pamb, tamb, mach_mca, MTO, nei)
dcx_oei = propu.oei_drag(aircraft, pamb, tamb)
cn_prop = propu.thrust_yaw_moment(aircraft, fn, pamb, mach_mca, dcx_oei)
x_cg = c_of_g.max_bwd_req_cg
cyb_vtp_req = (cn_prop * wing.mac) / ((xlc_vtp - x_cg) * aoa_max_vtp)
k_vtp_area = cyb_vtp_req / cyb_vtp
d_vtp_area = (k_vtp_area - 1) * vtp.area
return d_vtp_area