def eval_aerodynamics_design(aircraft):
"""
Defines high lift movable deflection settings
HLDconf varies from 0 (clean) to 1 (full deflected)
Typically : HLDconf = 1 ==> CzmaxLD
: HLDconf = 0.1 to 0.5 ==> CzmaxTO
"""
design_driver = aircraft.design_driver
wing = aircraft.wing
aerodynamics = aircraft.aerodynamics
mach = design_driver.cruise_mach
altp = design_driver.ref_cruise_altp
disa = 0
pamb, tamb, tstd, dtodz = earth.atmosphere(altp, disa)
aerodynamics.cruise_lod_max, aerodynamics.cz_cruise_lod_max = craft_aero.lod_max(
aircraft, pamb, tamb, mach)
aerodynamics.hld_conf_clean = 0 # By definition (0=<hld_conf=<1)
aerodynamics.cz_max_clean, Cz0 = craft_aero.high_lift(
wing, aerodynamics.hld_conf_clean)
aerodynamics.hld_conf_to = 0.3 # Take off (empirical setting)
aerodynamics.cz_max_to, Cz0 = craft_aero.high_lift(
wing, aerodynamics.hld_conf_to)
aerodynamics.hld_conf_ld = 1 # By definition (0=<hld_conf=<1), 1 is full landing
aerodynamics.cz_max_ld, Cz0 = craft_aero.high_lift(
wing, aerodynamics.hld_conf_ld)
return
def eval_wing_mass(aircraft):
"""
Wing mass & CG estimation
"""
weights = aircraft.weights
aerodynamics = aircraft.aerodynamics
wing = aircraft.wing
(cz_max_ld, cz0) = craft_aero.high_lift(wing, aerodynamics.hld_conf_ld)
A = 32 * wing.area**1.1
B = 4 * wing.span**2 * numpy.sqrt(weights.mtow * weights.mzfw)
C = 1.1e-6 * (1 + 2 * wing.aspect_ratio) / (1 + wing.aspect_ratio)
D = (0.6 * wing.t_o_c_r + 0.3 * wing.t_o_c_k +
0.1 * wing.t_o_c_t) * (wing.area / wing.span)
E = numpy.cos(wing.sweep)**2
F = 1200 * (cz_max_ld - 1.8)**1.5
wing.mass = A + (B * C) / (
D * E) + F # Shevell formula + high lift device regression
wing.c_g = 0.25*(wing.x_root + 0.40*wing.c_root) \
+ 0.55*(wing.x_kink + 0.40*wing.c_kink) \
+ 0.20*(wing.x_tip + 0.40*wing.c_tip)
return
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 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