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 fct_thrust(aircraft, state, rating, nei):
xg = state[state_dict["xg"]]
zg = state[state_dict["zg"]]
xg_d = state[state_dict["vgnd"]] * numpy.cos(state[state_dict["path"]])
zg_d = state[state_dict["vgnd"]] * numpy.sin(state[state_dict["path"]])
pamb, pamb_d, tamb, tamb_d, disa, disa_d, wx, wx_d, wz, wz_d = air(
xg, zg, xg_d, zg_d)
vsnd, vsnd_d = g_earth.sound_speed_grad(tamb, tamb_d)
vair = numpy.sqrt((xg_d - wx)**2 + (zg_d - wz)**2)
mach = vair / vsnd
fn, prop_data = propu.thrust(aircraft, pamb, tamb, mach, rating, nei)
return fn
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 = airplane_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 = 15.5 * ml_factor + 100.
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
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)
max_bwd_req_cg = xlc_vtp - (cn_prop * wing.mac) / (cyb_vtp * aoa_max_vtp)
c_of_g.max_bwd_oei_req_cg = c_of_g.max_bwd_req_cg
c_of_g.max_bwd_oei_cg = max_bwd_req_cg
c_of_g.max_bwd_oei_mass = tow
return
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)
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)
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