def req_oei_altp(propulsive_architecture):
if (propulsive_architecture < 4):
req_oei_altp_i = unit.m_ft(15000)
elif (propulsive_architecture == 4):
req_oei_altp_i = unit.m_ft(10000)
else:
raise Exception("propulsion.architecture index is out of range")
return req_oei_altp_i
def top_of_climb_altp(propulsive_architecture):
if (propulsive_architecture == 1):
top_of_climb_altp_i = unit.m_ft(35000)
elif (propulsive_architecture == 2):
top_of_climb_altp_i = unit.m_ft(35000)
elif (propulsive_architecture == 3):
top_of_climb_altp_i = unit.m_ft(35000)
elif (propulsive_architecture == 4):
top_of_climb_altp_i = unit.m_ft(20000)
else:
raise Exception("propulsion.architecture index is out of range")
return top_of_climb_altp_i
def handling_qualities_analysis(aircraft):
"""
Compute CG limits from handling qualities point of view
"""
# Forward limit : trim landing
#------------------------------------------------------------------------------------------------------
altp = unit.m_ft(0)
disa = 0
nei = 0
speed_mode = 1
hld_conf = aircraft.aerodynamics.hld_conf_ld
mass = aircraft.center_of_gravity.max_fwd_mass
cg_max_fwd_stall, speed, fn, aoa, ih, c_z, cx_trimmed = h_q.forward_cg_stall(
aircraft, altp, disa, nei, hld_conf, speed_mode, mass)
aircraft.center_of_gravity.max_fwd_trim_cg = cg_max_fwd_stall # Forward cg limit
# Backward limit : static stability
#------------------------------------------------------------------------------------------------------
stability_margin = 0.05
cg_max_bwd_stab = h_q.backward_cg_stab(aircraft, stability_margin)
aircraft.center_of_gravity.max_bwd_stab_cg = cg_max_bwd_stab # Backward cg limit
return
# Solve the geometric coupling between airframe and engines
#------------------------------------------------------------------------------------------------------
run.aircraft_pre_design(aircraft)
# Estimate all mass and CGs
#------------------------------------------------------------------------------------------------------
run.mass_mission_adaptation(aircraft)
#run.mass_estimation(aircraft)
# Calculate all airplane performances
#------------------------------------------------------------------------------------------------------
run.performance_analysis(aircraft)
# Print relevant output
#------------------------------------------------------------------------------------------------------
altp = unit.m_ft(35000)
disa = 0
pamb, tamb, tstd, dtodz = earth.atmosphere(altp, disa)
(MTO, MCN, MCL, MCR, FID) = aircraft.propulsion.rating_code
nei = 0
Fn, Data = propu.thrust(aircraft, pamb, tamb, cruise_mach, MTO, nei)
vsnd = earth.sound_speed(tamb)
tas = vsnd * cruise_mach
print("")
print("True air speed in cruise", "%.1f" % tas, " m/s")
print("Totalthrust in cruise", "%.0f" % Fn, " N")
print("")
print("Engine thrust = ",
"%.1f" % (aircraft.propulsion.reference_thrust_effective / 10), " daN")
print("Wing area = ", "%.1f" % aircraft.wing.area, " m2")
print('Criterion : CO2_metric = ', "%.4f" % (CO2_metric * 1000),
'kg/km/m0.48 (minimize)')
print('Criterion : block_CO2 = ', "%.0f" % block_CO2, 'kg, (minimize)')
print('Criterion : block_fuel = ', "%.2f" % block_fuel, 'kg, (minimize)')
print('Criterion : mtow_eff = ', "%.1f" % mtow_eff, 'kg, (minimize)')
print('Criterion : cash_op_cost = ', "%.2f" % cash_op_cost,
'$/trip, (minimize)')
print('Criterion : direct_op_cost = ', "%.2f" % direct_op_cost,
'$/trip, (minimize)')
print("-------------------------------------------")
print("Performance analysis : done")
# Handling_Qualities
#------------------------------------------------------------------------------------------------------
altp = unit.m_ft(0)
disa = 0
nei = 0
speed_mode = 1
hld_conf = ac.aerodynamics.hld_conf_ld
mass = ac.center_of_gravity.max_fwd_mass
cg_max_fwd_stall, speed, fn, aoa, ih, c_z, cx_trimmed = h_q.forward_cg_stall(
ac, altp, disa, nei, hld_conf, speed_mode, mass)
ac.center_of_gravity.max_fwd_trim_cg = cg_max_fwd_stall # Forward cg limit
stability_margin = 0.05
cg_max_bwd_stab = h_q.backward_cg_stab(ac, stability_margin)
def altp_app_speed():
altp_app_speed_i = unit.m_ft(0)
return altp_app_speed_i
def time_to_climb(aircraft, toc, disa, mass, vcas1, vcas2, mach):
"""
Time to climb to initial cruise altitude
"""
propulsion = aircraft.propulsion
(MTO, MCN, MCL, MCR, FID) = propulsion.rating_code
if (vcas1 > unit.mps_kt(250)):
print("time_to_climb_, vcas1 must be lower than or equal to 250kt")
if (vcas1 > vcas2):
print("time_to_climb_, vcas1 must be lower than or equal to vcas2")
cross_over_altp = earth.cross_over_altp(vcas2, mach)
if (cross_over_altp < unit.m_ft(1500)):
print("time_to_climb_, cross over altitude is too low")
if (toc < cross_over_altp):
cross_over_altp = toc
# Duration of initial climb
#-----------------------------------------------------------------------------------------------------------
altp = numpy.array([0., 0., 0.])
altp[0] = unit.m_ft(1500)
altp[2] = unit.m_ft(10000)
altp[1] = (altp[0] + altp[2]) / 2
nei = 0
speed_mode = 1 # Constant CAS
rating = MCL
v_z = numpy.array([0., 0., 0.])
[slope, v_z[0]] = flight.air_path(aircraft, nei, altp[0], disa, speed_mode,
vcas1, mass, rating)
[slope, v_z[1]] = flight.air_path(aircraft, nei, altp[1], disa, speed_mode,
vcas1, mass, rating)
[slope, v_z[2]] = flight.air_path(aircraft, nei, altp[2], disa, speed_mode,
vcas1, mass, rating)
if (numpy.extract(v_z < 0, v_z).size > 0):
print("time_to_climb_, Climb to acceleration altitude is not possible")
A = numpy.vander(altp, 3)
B = 1 / v_z
C = trinome(A, B)
time1 = ((C[0] * altp[2] / 3 + C[1] / 2) * altp[2] + C[2]) * altp[2]
time1 = time1 - (
(C[0] * altp[0] / 3 + C[1] / 2) * altp[0] + C[2]) * altp[0]
# Acceleration
#-----------------------------------------------------------------------------------------------------------
vcas = numpy.array([0., 0., 0.])
vcas[0] = vcas1
vcas[2] = vcas2
vcas[1] = (vcas[0] + vcas[2]) / 2
acc = numpy.array([0., 0., 0.])
acc[0] = flight.acceleration(aircraft, nei, altp[2], disa, speed_mode,
vcas[0], mass, rating)
acc[1] = flight.acceleration(aircraft, nei, altp[2], disa, speed_mode,
vcas[1], mass, rating)
acc[2] = flight.acceleration(aircraft, nei, altp[2], disa, speed_mode,
vcas[2], mass, rating)
if (numpy.extract(acc < 0, acc).size > 0):
print("time_to_climb_, acceleration is not possible")
A = numpy.vander(vcas, 3)
B = 1 / acc
C = trinome(A, B)
time2 = ((C[0] * vcas[2] / 3 + C[1] / 2) * vcas[2] + C[2]) * vcas[2]
time2 = time2 - (
(C[0] * vcas[0] / 3 + C[1] / 2) * vcas[0] + C[2]) * vcas[0]
# Duration of climb to cross over
#-----------------------------------------------------------------------------------------------------------
altp[0] = unit.m_ft(10000)
altp[2] = cross_over_altp
altp[1] = (altp[0] + altp[2]) / 2
[slope, v_z[0]] = flight.air_path(aircraft, nei, altp[0], disa, speed_mode,
vcas2, mass, rating)
[slope, v_z[1]] = flight.air_path(aircraft, nei, altp[1], disa, speed_mode,
vcas2, mass, rating)
[slope, v_z[2]] = flight.air_path(aircraft, nei, altp[2], disa, speed_mode,
vcas2, mass, rating)
if (numpy.extract(v_z < 0, v_z).size > 0):
print("time_to_climb_, Climb to cross over altitude is not possible")
A = numpy.vander(altp, 3)
B = 1 / v_z
C = trinome(A, B)
time3 = ((C[0] * altp[2] / 3 + C[1] / 2) * altp[2] + C[2]) * altp[2]
time3 = time3 - (
(C[0] * altp[0] / 3 + C[1] / 2) * altp[0] + C[2]) * altp[0]
# Duration of climb to altp
#-----------------------------------------------------------------------------------------------------------
if (cross_over_altp < toc):
altp[0] = cross_over_altp
altp[2] = toc
altp[1] = (altp[0] + altp[2]) / 2
speed_mode = 2 # mach
[slope, v_z[0]] = flight.air_path(aircraft, nei, altp[0], disa,
speed_mode, mach, mass, rating)
[slope, v_z[1]] = flight.air_path(aircraft, nei, altp[1], disa,
speed_mode, mach, mass, rating)
[slope, v_z[2]] = flight.air_path(aircraft, nei, altp[2], disa,
speed_mode, mach, mass, rating)
if (numpy.extract(v_z < 0, v_z).size > 0):
print("time_to_climb_, Climb to top of climb is not possible")
A = numpy.vander(altp, 3)
B = 1 / v_z
C = trinome(A, B)
time4 = ((C[0]*altp[2]/3 + C[1]/2)*altp[2] + C[2])*altp[2] \
- ((C[0]*altp[0]/3 + C[1]/2)*altp[0] + C[2])*altp[0]
else:
time4 = 0
# Total time
#-----------------------------------------------------------------------------------------------------------
ttc = time1 + time2 + time3 + time4
return ttc
