def speed_from_lift(aircraft, pamb, cz, mass):
wing = aircraft.wing
g = earth.gravity()
gam = earth.heat_ratio()
mach = numpy.sqrt((mass * g) / (0.5 * gam * pamb * wing.area * cz))
return mach
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 lift_from_speed(aircraft, pamb, mach, mass):
wing = aircraft.wing
g = earth.gravity()
gam = earth.heat_ratio()
c_z = (2 * mass * g) / (gam * pamb * mach**2 * wing.area)
return c_z
def thrust_yaw_moment(aircraft, fn, pamb, mach, dcx_oei):
"""
Assumed right engine inoperative
"""
propulsion = aircraft.propulsion
wing = aircraft.wing
gam = earth.heat_ratio()
if (propulsion.architecture == 1):
nacelle = aircraft.turbofan_nacelle
elif (propulsion.architecture == 2):
nacelle = aircraft.turbofan_nacelle
elif (propulsion.architecture == 3):
nacelle = aircraft.turbofan_nacelle
elif (propulsion.architecture == 4):
nacelle = aircraft.turboprop_nacelle
else:
raise Exception("propulsion.architecture index is out of range")
cn_prop = (nacelle.y_ext /
wing.mac) * (fn /
(0.5 * gam * pamb * mach**2 * wing.area) + dcx_oei)
return cn_prop
def thrust_pitch_moment(aircraft, fn, pamb, mach, dcx_oei):
propulsion = aircraft.propulsion
wing = aircraft.wing
gam = earth.heat_ratio()
if (propulsion.architecture == 1):
nacelle = aircraft.turbofan_nacelle
elif (propulsion.architecture == 2):
nacelle = aircraft.turbofan_nacelle
elif (propulsion.architecture == 3):
nacelle = aircraft.turbofan_nacelle
elif (propulsion.architecture == 4):
nacelle = aircraft.turboprop_nacelle
else:
raise Exception("propulsion.architecture index is out of range")
cm_prop = nacelle.z_ext * (dcx_oei - fn /
(0.5 * gam * pamb * mach**2 * wing.area))
return cm_prop
def forward_cg_stall(aircraft, altp, disa, nei, hld_conf, speed_mode, mass):
"""
Computes max forward trimmable CG position at stall speed
"""
wing = aircraft.wing
htp = aircraft.horizontal_tail
gam = earth.heat_ratio()
[pamb, tamb, tstd, dtodz] = earth.atmosphere(altp, disa)
[cz_max_wing, cz0] = craft_aero.high_lift(
wing, hld_conf) # Wing maximum lift coefficient without margin
[cza_htp, xlc_htp, aoa_max_htp,
ki_htp] = frame_aero.htp_aero_data(aircraft)
cz_max_htp = cza_htp * aoa_max_htp
c_z = cz_max_wing - cz_max_htp # Max forward Cg assumed, HTP has down lift
mach = flight.speed_from_lift(aircraft, pamb, c_z, mass)
[cza_wo_htp, xlc_wo_htp,
ki_wing] = frame_aero.wing_aero_data(aircraft, mach, hld_conf)
if (nei > 0):
dcx_oei = nei * propu.oei_drag(pamb, mach)
else:
dcx_oei = 0
dw_angle = frame_aero.wing_downwash(
aircraft, cz_max_wing) # Downwash angle due to the wing
cx_basic, lod_trash = craft_aero.drag(
aircraft, pamb, tamb, mach,
cz_max_wing) # By definition of the drag_ function
cxi_htp = (ki_htp * cz_max_htp**2) * (htp.area / wing.area
) # Induced drag generated by HTP
cx_inter = cz_max_htp * dw_angle # Interaction drag (due to downwash)
cx_trimmed = cx_basic + cxi_htp + cx_inter + dcx_oei
fn = 0.5 * gam * pamb * mach**2 * wing.area * cx_trimmed
cm_prop = propu.thrust_pitch_moment(aircraft, fn, pamb, mach, dcx_oei)
cg_max_fwd_stall = (cm_prop + xlc_wo_htp * cz_max_wing -
xlc_htp * cz_max_htp) / (cz_max_wing - cz_max_htp)
aoa_wing = (cz_max_wing - cz0) / cza_wo_htp # Wing angle of attack
aoa = aoa_wing - wing.setting # Reference angle of attack (fuselage axis versus air speed)
ih = -aoa + dw_angle - aoa_max_htp # HTP trim setting
speed = flight.get_speed(pamb, speed_mode, mach)
return cg_max_fwd_stall, speed, fn, aoa, ih, c_z, cx_trimmed
def corrected_air_flow(Ptot, Ttot, Mach):
"""
Computes the corrected air flow per square meter
"""
R = earth.gaz_constant()
gam = earth.heat_ratio()
f_M = Mach * (1. + 0.5 * (gam - 1) * Mach**2)**(-(gam + 1.) / (2. *
(gam - 1.)))
CQoA = (math.sqrt(gam / R) * Ptot / math.sqrt(Ttot)) * f_M
return CQoA
def eval_pre_design_wing(aircraft):
"""
Wing predesign
"""
design_driver = aircraft.design_driver
fuselage = aircraft.fuselage
vtp = aircraft.vertical_tail
nacelle = aircraft.turbofan_nacelle
weights = aircraft.weights
wing = aircraft.wing
wing.t_o_c_t = 0.10
wing.t_o_c_k = wing.t_o_c_t + 0.01
wing.t_o_c_r = wing.t_o_c_k + 0.03
wing.sweep = 1.6 * max(0,
(design_driver.cruise_mach - 0.5)) # Empirical law
wing.dihedral = unit.rad_deg(5)
if (wing.morphing == 1): # Aspect ratio is driving parameter
wing.span = numpy.sqrt(wing.aspect_ratio * wing.area)
elif (wing.morphing == 2): # Span is driving parameter
wing.aspect_ratio = wing.span**2 / wing.area
else:
print("geometry_predesign_, wing_morphing index is unkown")
# Correlation between span loading and tapper ratio
wing.taper_ratio = 0.3 - 0.025 * (1e-3 * weights.mtow / wing.span)
# Factor 1.64 accounts for the part of HTP ref area hidden in the fuselage
wing.net_wetted_area = 1.64 * wing.area
wing.y_kink = nacelle.y_ext
wing.y_root = 0.5 * fuselage.width
wing.y_tip = 0.5 * wing.span
if (15 < unit.deg_rad(wing.sweep)): # With kink
Phi100intTE = max(0, 2 * (wing.sweep - unit.rad_deg(32)))
tan_phi100 = numpy.tan(Phi100intTE)
A = ((1 - 0.25 * wing.taper_ratio) * wing.y_kink +
0.25 * wing.taper_ratio * wing.y_root - wing.y_tip) / (
0.75 * wing.y_kink + 0.25 * wing.y_root - wing.y_tip)
B = (numpy.tan(wing.sweep) - tan_phi100) * (
(wing.y_tip - wing.y_kink) *
(wing.y_kink - wing.y_root)) / (0.25 * wing.y_root +
0.75 * wing.y_kink - wing.y_tip)
wing.c_root = (wing.area - B * (wing.y_tip - wing.y_root)) / (
wing.y_root + wing.y_kink + A *
(wing.y_tip - wing.y_root) + wing.taper_ratio *
(wing.y_tip - wing.y_kink))
wing.c_kink = A * wing.c_root + B
wing.c_tip = wing.taper_ratio * wing.c_root
else: # Without kink
wing.c_root = 2 * wing.area / (
2 * wing.y_root * (1 - wing.taper_ratio) +
(1 + wing.taper_ratio) * numpy.sqrt(wing.aspect_ratio * wing.area))
wing.c_tip = wing.taper_ratio * wing.c_root
wing.c_kink = ((wing.y_tip - wing.y_kink) * wing.c_root +
(wing.y_kink - wing.y_root) * wing.c_tip) / (
wing.y_tip - wing.y_root)
tan_phi0 = 0.25 * (wing.c_kink - wing.c_tip) / (
wing.y_tip - wing.y_kink) + numpy.tan(wing.sweep)
wing.mac = 2.*(3*wing.y_root*wing.c_root**2 \
+(wing.y_kink-wing.y_root)*(wing.c_root**2+wing.c_kink**2+wing.c_root*wing.c_kink) \
+(wing.y_tip-wing.y_kink)*(wing.c_kink**2+wing.c_tip**2+wing.c_kink*wing.c_tip) \
)/(3*wing.area)
wing.y_mac = ( 3*wing.c_root*wing.y_root**2 \
+(wing.y_kink-wing.y_root)*(wing.c_kink*(wing.y_root+wing.y_kink*2.)+wing.c_root*(wing.y_kink+wing.y_root*2.)) \
+(wing.y_tip-wing.y_kink)*(wing.c_tip*(wing.y_kink+wing.y_tip*2.)+wing.c_kink*(wing.y_tip+wing.y_kink*2.)) \
)/(3*wing.area)
x_mac_local = ( (wing.y_kink-wing.y_root)*tan_phi0*((wing.y_kink-wing.y_root)*(wing.c_kink*2.+wing.c_root) \
+(wing.y_tip-wing.y_kink)*(wing.c_kink*2.+wing.c_tip))+(wing.y_tip-wing.y_root)*tan_phi0*(wing.y_tip-wing.y_kink)*(wing.c_tip*2.+wing.c_kink) \
)/(3*wing.area)
wing.x_root = vtp.x_mac + 0.25 * vtp.mac - vtp.lever_arm - 0.25 * wing.mac - x_mac_local
wing.x_kink = wing.x_root + (wing.y_kink - wing.y_root) * tan_phi0
wing.x_tip = wing.x_root + (wing.y_tip - wing.y_root) * tan_phi0
wing.x_mac = wing.x_root+( (wing.x_kink-wing.x_root)*((wing.y_kink-wing.y_root)*(wing.c_kink*2.+wing.c_root) \
+(wing.y_tip-wing.y_kink)*(wing.c_kink*2.+wing.c_tip))+(wing.x_tip-wing.x_root)*(wing.y_tip-wing.y_kink)*(wing.c_tip*2.+wing.c_kink) \
)/(wing.area*3.)
if (wing.attachment == 1):
wing.z_root = 0
else:
wing.z_root = fuselage.height - 0.5 * wing.t_o_c_r * wing.c_root
wing.z_kink = wing.z_root + (wing.y_kink - wing.y_root) * numpy.tan(
wing.dihedral)
wing.z_tip = wing.z_root + (wing.y_tip - wing.y_root) * numpy.tan(
wing.dihedral)
# Wing setting
#-----------------------------------------------------------------------------------------------------------
g = earth.gravity()
gam = earth.heat_ratio()
disa = 0
rca = design_driver.ref_cruise_altp
mach = design_driver.cruise_mach
mass = 0.95 * weights.mtow
pamb, tamb, tstd, dtodz = earth.atmosphere(rca, disa)
cza_wo_htp = frame_aero.cza_wo_htp(mach, fuselage.width, wing.aspect_ratio,
wing.span, wing.sweep)
# AoA = 2.5° at cruise start
wing.setting = (0.97 * mass * g) / (0.5 * gam * pamb * mach**2 * wing.area
* cza_wo_htp) - unit.rad_deg(2.5)
return
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_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)