def nominal_mission(aircraft):
"""
Compute nominal mission with range as input
"""
disa = 0
altp = aircraft.design_driver.ref_cruise_altp
mach = aircraft.design_driver.cruise_mach
nei = 0
aircraft.nominal_mission.payload = aircraft.payload.nominal
aircraft.nominal_mission.range = aircraft.design_driver.design_range
aircraft.nominal_mission.tow = aircraft.weights.mtow
(MTO, MCN, MCL, MCR, FID) = aircraft.propulsion.rating_code
pamb, tamb, tstd, dtodz = earth.atmosphere(altp, disa)
lod_max, _ = craft_aero.lod_max(aircraft, pamb, tamb, mach)
sfc = propu.sfc(aircraft, pamb, tamb, mach, MCR, nei)
aircraft.high_speed.cruise_lod = lod_max
aircraft.high_speed.cruise_sfc = sfc
payload = aircraft.nominal_mission.payload
range = aircraft.nominal_mission.range
tow = aircraft.nominal_mission.tow
block_fuel, block_time, total_fuel = perfo.mission(aircraft, range, tow,
altp, mach, disa)
aircraft.nominal_mission.block_fuel = block_fuel
aircraft.nominal_mission.block_time = block_time
aircraft.nominal_mission.total_fuel = total_fuel
return
def specific_air_range(aircraft, altp, mass, mach, disa):
propulsion = aircraft.propulsion
(MTO, MCN, MCL, MCR, FID) = propulsion.rating_code
g = earth.gravity()
pamb, tamb, tstd, dtodz = earth.atmosphere(altp, disa)
vsnd = earth.sound_speed(tamb)
Cz = flight.lift_from_speed(aircraft, pamb, mach, mass)
[Cx, LoD] = craft_aero.drag(aircraft, pamb, tamb, mach, Cz)
nei = 0.
sfc = propu.sfc(aircraft, pamb, tamb, mach, MCR, nei)
sar = (vsnd * mach * LoD) / (mass * g * sfc)
return sar
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
print('Coupling : delta_mlw = ', "%.2f" % delta_mlw, 'kg, (= 0 ?)')
print("-------------------------------------------")
print("Mass and CG estimation : done")
# Nominal mission
#------------------------------------------------------------------------------------------------------
disa = 0
altp = ac.design_driver.ref_cruise_altp
mach = ac.design_driver.cruise_mach
nei = 0
(MTO, MCN, MCL, MCR, FID) = ac.propulsion.rating_code
pamb, tamb, tstd, dtodz = earth.atmosphere(altp, disa)
lod_max, _ = craft_aero.lod_max(ac, pamb, tamb, mach)
sfc = propu.sfc(ac, pamb, tamb, mach, MCR, nei)
ac.high_speed.cruise_lod = lod_max
ac.high_speed.cruise_sfc = sfc
ac.nominal_mission.payload = ac.payload.nominal
ac.nominal_mission.range = ac.design_driver.design_range
ac.nominal_mission.tow = ac.weights.mtow
payload = ac.nominal_mission.payload
range = ac.nominal_mission.range
tow = ac.nominal_mission.tow
block_fuel, block_time, total_fuel = perfo.mission(ac, range, tow, altp, mach,
disa)
def performance_analysis(aircraft):
"""
Compute operational performances
"""
# Nominal mission (here : range is an output)
#------------------------------------------------------------------------------------------------------
disa = 0.
altp = aircraft.design_driver.ref_cruise_altp
mach = aircraft.design_driver.cruise_mach
nei = 0.
(MTO, MCN, MCL, MCR, FID) = aircraft.propulsion.rating_code
pamb, tamb, tstd, dtodz = earth.atmosphere(altp, disa)
lod_max, _ = craft_aero.lod_max(aircraft, pamb, tamb, mach)
sfc = propu.sfc(aircraft, pamb, tamb, mach, MCR, nei)
aircraft.high_speed.cruise_lod = lod_max
aircraft.high_speed.cruise_sfc = sfc
aircraft.nominal_mission.payload = aircraft.payload.nominal
aircraft.nominal_mission.range = aircraft.design_driver.design_range
aircraft.nominal_mission.tow = aircraft.weights.mtow
payload = aircraft.nominal_mission.payload
tow = aircraft.nominal_mission.tow
range, block_fuel, block_time, total_fuel = sub_proc.mission_range(
aircraft, tow, payload, altp, mach, disa)
aircraft.nominal_mission.range = range
aircraft.nominal_mission.block_fuel = block_fuel
aircraft.nominal_mission.block_time = block_time
aircraft.nominal_mission.total_fuel = total_fuel
# Ceilings
#------------------------------------------------------------------------------------------------------
toc = aircraft.design_driver.top_of_climb_altp
oei_ceil_req = aircraft.low_speed.req_oei_altp
vz_clb, vz_crz, oei_path, oei_mach = perfo.ceilings(
aircraft, toc, oei_ceil_req)
aircraft.high_speed.eff_vz_climb = vz_clb
aircraft.high_speed.eff_vz_cruise = vz_crz
aircraft.low_speed.eff_oei_path = oei_path
# Time to climb to requested altitude
#------------------------------------------------------------------------------------------------------
toc = aircraft.high_speed.req_toc_altp
disa = 0
mass = aircraft.weights.mtow
vcas1 = aircraft.high_speed.cas1_ttc
vcas2 = aircraft.high_speed.cas2_ttc
mach = aircraft.design_driver.cruise_mach
ttc = perfo.time_to_climb(aircraft, toc, disa, mass, vcas1, vcas2, mach)
aircraft.high_speed.eff_ttc = ttc
# Take off field length
#------------------------------------------------------------------------------------------------------
altp = aircraft.low_speed.altp_tofl
disa = aircraft.low_speed.disa_tofl
mass = aircraft.weights.mtow
hld_conf_to = aircraft.aerodynamics.hld_conf_to
tofl, seg2_path, eff_kvs1g, limitation = perfo.take_off_field_length(
aircraft, altp, disa, mass, hld_conf_to)
aircraft.low_speed.eff_tofl = tofl
aircraft.low_speed.eff_kvs1g = eff_kvs1g
aircraft.low_speed.seg2_path = seg2_path
aircraft.low_speed.limitation = limitation
# Approach speed
#------------------------------------------------------------------------------------------------------
altp = aircraft.low_speed.altp_app_speed
disa = aircraft.low_speed.disa_app_speed
mass = aircraft.weights.mlw
hld_conf_ld = aircraft.aerodynamics.hld_conf_to
app_speed = perfo.approach_speed(aircraft, altp, disa, mass, hld_conf_ld)
aircraft.low_speed.eff_app_speed = app_speed
# Environment
#------------------------------------------------------------------------------------------------------
CO2_metric, rgf = environ.fuel_efficiency_metric(aircraft)
aircraft.environmental_impact.rgf = rgf
aircraft.environmental_impact.CO2_metric = CO2_metric
# Cost mission
#-----------------------------------------------------------------------------------------------------------------------------------------------
altp = aircraft.design_driver.ref_cruise_altp
mach = aircraft.design_driver.cruise_mach
disa = aircraft.cost_mission.disa
range = aircraft.cost_mission.range
payload = aircraft.payload.nominal
aircraft.cost_mission.payload = payload
tow, block_fuel, block_time, total_fuel = sub_proc.mission_tow(
aircraft, payload, range, altp, mach, disa)
aircraft.cost_mission.block_fuel = block_fuel
aircraft.cost_mission.block_time = block_time
aircraft.cost_mission.total_fuel = total_fuel
aircraft.cost_mission.block_CO2 = block_fuel * aircraft.environmental_impact.CO2_index
# Economics
#------------------------------------------------------------------------------------------------------
direct_op_cost, cash_op_cost, block_fuel, engine_price, airplane_price = costing.operating_costs(
aircraft, block_fuel, block_time)
aircraft.economics.engine_price = engine_price
aircraft.economics.airplane_price = airplane_price
aircraft.economics.direct_operating_cost = direct_op_cost
aircraft.economics.cash_operating_cost = cash_op_cost
return
def mission(aircraft, dist_range, tow, altp, mach, disa):
"""
Mission computation using breguet equation, fixed L/D and fixed sfc
"""
engine = aircraft.turbofan_engine
propulsion = aircraft.propulsion
battery = aircraft.battery
(MTO, MCN, MCL, MCR, FID) = propulsion.rating_code
g = earth.gravity()
pamb, tamb, tstd, dtodz = earth.atmosphere(altp, disa)
vsnd = earth.sound_speed(tamb)
tas = vsnd * mach
lod_max, cz_lod_max = craft_aero.lod_max(aircraft, pamb, tamb, mach)
lod_cruise = 0.95 * lod_max
nei = 0.
sfc = propu.sfc(aircraft, pamb, tamb, mach, MCR, nei)
if (propulsion.architecture == 2):
fn, sec, data = propu.hybrid_thrust(aircraft, pamb, tamb, mach, MCR,
nei)
if (propulsion.architecture == 3):
fn, sec, data = propu.hybrid_thrust(aircraft, pamb, tamb, mach, MCR,
nei)
# Departure ground phases
#-----------------------------------------------------------------------------------------------------------
fuel_taxi_out = (34 + 2.3e-4 * engine.reference_thrust) * engine.n_engine
time_taxi_out = 540
fuel_take_off = 1e-4 * (2.8 + 2.3 / engine.bpr) * tow
time_take_off = 220 * tow / (engine.reference_thrust * engine.n_engine)
# Mission leg
#-----------------------------------------------------------------------------------------------------------
if (propulsion.architecture == 1):
fuel_mission = tow * (1 - numpy.exp(-(sfc * g * dist_range) /
(tas * lod_cruise)))
elif (propulsion.architecture == 2):
fuel_mission = tow*(1-numpy.exp(-(sfc*g*dist_range)/(tas*lod_cruise))) \
- (sfc/sec)*battery.energy_cruise
elif (propulsion.architecture == 3):
fuel_mission = tow*(1-numpy.exp(-(sfc*g*dist_range)/(tas*lod_cruise))) \
- (sfc/sec)*battery.energy_cruise
elif (propulsion.architecture == 4):
fuel_mission = tow * (1 - numpy.exp(-(sfc * g * dist_range) /
(tas * lod_cruise)))
else:
raise Exception("propulsion.architecture index is out of range")
time_mission = 1.09 * (dist_range / tas)
l_w = tow - fuel_mission
# Arrival ground phases
#-----------------------------------------------------------------------------------------------------------
fuel_landing = 1e-4 * (0.5 + 2.3 / engine.bpr) * l_w
time_landing = 180
fuel_taxi_in = (26 + 1.8e-4 * engine.reference_thrust) * engine.n_engine
time_taxi_in = 420
# Block fuel and time
#-----------------------------------------------------------------------------------------------------------
block_fuel = fuel_taxi_out + fuel_take_off + fuel_mission + fuel_landing + fuel_taxi_in
time_block = time_taxi_out + time_take_off + time_mission + time_landing + time_taxi_in
# Diversion and holding reserve fuel
#-----------------------------------------------------------------------------------------------------------
fuel_diversion = l_w * (1 -
numpy.exp(-(sfc * g * regul.diversion_range()) /
(tas * lod_cruise)))
fuel_holding = sfc * (l_w * g / lod_max) * regul.holding_time()
# Total
#-----------------------------------------------------------------------------------------------------------
fuel_total = fuel_mission * (
1 + regul.reserve_fuel_ratio()) + fuel_diversion + fuel_holding
#-----------------------------------------------------------------------------------------------------------
return block_fuel, time_block, fuel_total