def turbofan_thrust(aircraft, Pamb, Tamb, Mach, rating, nei):
"""
Calculation of thrust for pure turbofan airplane
Warning : ALL engine thrust returned
"""
engine = aircraft.turbofan_engine
nacelle = aircraft.turbofan_nacelle
factor = engine.rating_factor # [MTO,MCN,MCL,MCR,FID]
kth = 0.475*Mach**2 + 0.091*(engine.bpr/10)**2 \
- 0.283*Mach*engine.bpr/10 \
- 0.633*Mach - 0.081*engine.bpr/10 + 1.192
(rho, sig) = earth.air_density(Pamb, Tamb)
fn0 = factor[rating] * kth * engine.reference_thrust * sig**0.75
fn_core = fn0 * engine.core_thrust_ratio # Core thrust
fn_fan0 = fn0 * (1 - engine.core_thrust_ratio) # Fan thrust
Vsnd = earth.sound_speed(Tamb)
Vair = Vsnd * Mach
shaft_power0 = fn_fan0 * Vair / nacelle.efficiency_prop # Available total shaft power for one engine
fn = fn0 * (engine.n_engine - nei) # All turbofan thrust
data = (fn_core, fn_fan0, fn0, shaft_power0
) # Data for ONE turbofan engine
return fn, data
def approach_speed(aircraft, altp, disa, mass, hld_conf):
"""
Minimum approach speed (VLS)
"""
wing = aircraft.wing
g = earth.gravity()
czmax, trash = craft_aero.high_lift(wing, hld_conf)
stall_margin = regul.kvs1g_min_landing()
[pamb, tamb, tstd, dtodz] = earth.atmosphere(altp, disa)
[rho, sig] = earth.air_density(pamb, tamb)
vapp = numpy.sqrt(
(mass * g) / (0.5 * rho * wing.area * (czmax / stall_margin**2)))
return vapp
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
def eval_bli_nacelle_design(this_nacelle, Pamb, Tamb, Mach, shaft_power,
hub_width, body_length, body_width):
"""
BLI nacelle design
"""
gam = earth.heat_ratio()
r = earth.gaz_constant()
Cp = earth.heat_constant(gam, r)
(rho, sig) = earth.air_density(Pamb, Tamb)
Vsnd = earth.sound_speed(Tamb)
Re = craft_aero.reynolds_number(Pamb, Tamb, Mach)
Vair = Vsnd * Mach
# Precalculation of the relation between d0 and d1
#-----------------------------------------------------------------------------------------------------------
body_bnd_layer = eval_boundary_layer(body_width, hub_width)
# Electrical nacelle geometry : e-nacelle diameter is size by cruise conditions
#-----------------------------------------------------------------------------------------------------------
r0 = 0.5 * body_width # Radius of the fuselage, supposed constant
d0 = craft_aero.boundary_layer(
Re, body_length
) # theoritical thickness of the boundary layer without taking account of fuselage tapering
r1 = 0.5 * hub_width # Radius of the hub of the efan nacelle
d1 = lin_interp_1d(d0, body_bnd_layer[:, 0],
body_bnd_layer[:,
1]) # Thickness of the BL around the hub
deltaV = 2. * Vair * (
this_nacelle.efficiency_fan / this_nacelle.efficiency_prop - 1.
) # speed variation produced by the fan
PwInput = this_nacelle.efficiency_fan * shaft_power # kinetic energy produced by the fan
#===========================================================================================================
def fct_power_1(y, PwInput, deltaV, rho, Vair, r1, d1):
(q0, q1, q2, v1, dVbli) = craft_aero.air_flows(rho, Vair, r1, d1, y)
Vinlet = Vair - dVbli
Vjet = Vinlet + deltaV
Pw = 0.5 * q1 * (Vjet**2 - Vinlet**2)
y = PwInput - Pw
return y
#-----------------------------------------------------------------------------------------------------------
fct_arg = (PwInput, deltaV, rho, Vair, r1, d1)
# Computation of y1 : thickness of the vein swallowed by the inlet
output_dict = fsolve(fct_power_1, x0=d1, args=fct_arg, full_output=True)
y1 = output_dict[0][0]
(q0, q1, q2, v1, dVbli) = craft_aero.air_flows(rho, Vair, r1, d1, y1)
MachInlet = v1 / Vsnd # Mean Mach number at inlet position
Ptot = earth.total_pressure(
Pamb, MachInlet) # Stagnation pressure at inlet position
Ttot = earth.total_temperature(
Tamb, MachInlet) # Stagnation temperature at inlet position
MachFan = 0.5 # required Mach number at fan position
CQoA1 = craft_aero.corrected_air_flow(
Ptot, Ttot, MachFan) # Corrected air flow per area at fan position
eFanArea = q1 / CQoA1 # Fan area around the hub
fan_width = math.sqrt(hub_width**2 +
4 * eFanArea / math.pi) # Fan diameter
Vjet = v1 + deltaV # Jet velocity
TtotJet = Ttot + shaft_power / (
q1 * Cp) # Stagnation pressure increases due to introduced work
Tstat = TtotJet - 0.5 * Vjet**2 / Cp # static temperature
VsndJet = math.sqrt(gam * r * Tstat) # Sound velocity at nozzle exhaust
MachJet = Vjet / VsndJet # Mach number at nozzle output
PtotJet = earth.total_pressure(
Pamb, MachJet) # total pressure at nozzle exhaust (P = Pamb)
CQoA2 = craft_aero.corrected_air_flow(
PtotJet, TtotJet,
MachJet) # Corrected air flow per area at nozzle output
nozzle_area = q1 / CQoA2 # Fan area around the hub
nozzle_width = math.sqrt(4 * nozzle_area / math.pi) # Nozzle diameter
this_nacelle.hub_width = hub_width
this_nacelle.fan_width = fan_width
this_nacelle.nozzle_width = nozzle_width
this_nacelle.nozzle_area = nozzle_area
this_nacelle.width = 1.19 * fan_width # Surrounding structure
this_nacelle.length = 1.68 * this_nacelle.width
this_nacelle.net_wetted_area = math.pi * this_nacelle.width * this_nacelle.length # Nacelle wetted area
this_nacelle.bnd_layer = body_bnd_layer
this_nacelle.body_length = body_length
return
def fan_thrust(nacelle, Pamb, Tamb, Mach, Vair, PwShaft):
"""
Compute the thrust of a fan of given geometry swallowing
the boundary layer (BL) of a body of given geometry
The amount of swallowed BL depends on the given shaft power
and flying conditions
"""
bnd_layer = nacelle.bnd_layer
Re = craft_aero.reynolds_number(Pamb, Tamb, Mach)
(rho, sig) = earth.air_density(Pamb, Tamb)
Vsnd = earth.sound_speed(Tamb)
d0 = craft_aero.boundary_layer(
Re, nacelle.body_length
) #Theoritical thickness of the boundary layer without taking account of fuselage tapering
d1 = lin_interp_1d(d0, bnd_layer[:, 0],
bnd_layer[:, 1]) # Using the precomputed relation
r1 = 0.5 * nacelle.hub_width # Radius of the hub of the eFan nacelle
PwInput = nacelle.efficiency_fan * PwShaft
#===========================================================================================================
def fct_power(q, PwInput, Vair, Vsnd, eNozzleArea):
Vinlet = Vair
MachInlet = Vinlet / Vsnd # Mean Mach number at inlet position
Ptot = earth.total_pressure(
Pamb, MachInlet) #total pressure at inlet position
Ttot = earth.total_temperature(
Tamb, MachInlet) # Total temperature at inlet position
Vjet = math.sqrt(2. * PwInput / q + Vinlet**2)
MachJet = Vjet / Vsnd
CQoA1 = craft_aero.corrected_air_flow(
Ptot, Ttot, MachJet) # Corrected air flow per area at fan position
q0 = CQoA1 * eNozzleArea
y = q - q0
return y
#-----------------------------------------------------------------------------------------------------------
eNozzleArea = nacelle.nozzle_area
fct_arg = (PwInput, Vair, Vsnd, eNozzleArea)
(q0init, q1, q2, V1, dV) = craft_aero.air_flows(rho, Vair, r1, d1, 1.00)
# Computation of y1 : thikness of the vein swallowed by the inlet
output_dict = fsolve(fct_power, x0=q0init, args=fct_arg, full_output=True)
q0 = output_dict[0][0]
Vinlet = Vair
Vjet = math.sqrt(2. * PwInput / q0 + Vinlet**2)
eFn = q0 * (Vjet - Vinlet)
return (eFn, q0)
def fan_thrust_with_bli(nacelle, Pamb, Tamb, Mach, Vair, PwShaft):
"""
Compute the thrust of a fan of a given geometry swallowing
the boundary layer (BL) of a body of a given geometry
The amount of swallowed BL depends on the given shaft power and flying
conditions.
"""
bnd_layer = nacelle.bnd_layer
gam = earth.heat_ratio()
r = earth.gaz_constant()
Cp = earth.heat_constant(gam, r)
Re = craft_aero.reynolds_number(Pamb, Tamb, Mach)
(rho, sig) = earth.air_density(Pamb, Tamb)
Vsnd = earth.sound_speed(Tamb)
d0 = craft_aero.boundary_layer(
Re, nacelle.body_length
) # theorical thickness of the boundary layer without taking account of fuselage tapering
r1 = 0.5 * nacelle.hub_width # Radius of the hub of the eFan nacelle
d1 = lin_interp_1d(d0, bnd_layer[:, 0],
bnd_layer[:, 1]) # Using the precomputed relation
#===========================================================================================================
def fct_power_bli(y, PwShaft, Tamb, Pamb, rho, Mach, Vair, Vsnd, r1, d1,
nozzle_area):
Ttot = earth.total_temperature(
Tamb, Mach) # Stagnation temperature at inlet position
(q0, q1, q2, Vinlet,
dVbli) = craft_aero.air_flows(rho, Vair, r1, d1, y)
Tstat = Ttot - 0.5 * Vinlet**2 / Cp # Static temperature at inlet position
Vsnd_inlet = earth.sound_speed(Tstat) # Sound speed at inlet position
MachInlet = Vinlet / Vsnd_inlet # Mean Mach number at inlet position
PwInput = nacelle.efficiency_fan * PwShaft
Vjet = math.sqrt(2. * PwInput / q1 + Vinlet**2)
TtotJet = Ttot + PwShaft / (
q1 * Cp) # Stagnation temperature increases due to introduced work
Tstat = TtotJet - 0.5 * Vjet**2 / Cp # Static temperature
VsndJet = earth.sound_speed(Tstat) # Sound speed at nozzle exhaust
MachJet = Vjet / VsndJet # Mach number at nozzle output
PtotJet = earth.total_pressure(
Pamb, MachJet) # total pressure at nozzle exhaust (P = Pamb)
CQoA1 = craft_aero.corrected_air_flow(
PtotJet, TtotJet,
MachJet) # Corrected air flow per area at fan position
q = CQoA1 * nozzle_area
y = q1 - q
return y
#-----------------------------------------------------------------------------------------------------------
nozzle_area = nacelle.nozzle_area
fct_arg = (PwShaft, Tamb, Pamb, rho, Mach, Vair, Vsnd, r1, d1, nozzle_area)
# Computation of y1 : thikness of the vein swallowed by the inlet
output_dict = fsolve(fct_power_bli,
x0=0.50,
args=fct_arg,
full_output=True)
y = output_dict[0][0]
Ttot = earth.total_temperature(
Tamb, Mach) # Stagnation temperature at inlet position
(q0, q1, q2, Vinlet, dVbli) = craft_aero.air_flows(rho, Vair, r1, d1, y)
Tstat = Ttot - 0.5 * Vinlet**2 / Cp # Static temperature at inlet position
Vsnd_inlet = earth.sound_speed(Tstat) # Sound speed at inlet position
MachInlet = Vinlet / Vsnd_inlet # Mean Mach number at inlet position
PwInput = nacelle.efficiency_fan * PwShaft
Vjet = math.sqrt(2. * PwInput / q1 + Vinlet**2)
eFn = q1 * (Vjet - Vinlet)
return (eFn, q1, dVbli)