Source code for RCAIDE.Library.Methods.Aerodynamics.Common.Lift.compute_max_lift_coeff
# ----------------------------------------------------------------------
# Imports
# ----------------------------------------------------------------------
#RCAIDE Imports
from RCAIDE import *
from RCAIDE.Framework.Core import Units, Data
from RCAIDE.Library.Components import Wings
from RCAIDE.Library.Methods.Aerodynamics.Common.Lift.compute_slat_lift import compute_slat_lift
from RCAIDE.Library.Methods.Aerodynamics.Common.Lift.compute_flap_lift import compute_flap_lift
# ----------------------------------------------------------------------
# compute_max_lift_coeff
# ----------------------------------------------------------------------
[docs]
def compute_max_lift_coeff(state,settings,geometry):
"""Computes the maximum lift coefficient associated with an aircraft high lift system
Assumptions:
None
Source:
Unknown
Inputs:
analyses.max_lift_coefficient_factor [Unitless]
vehicle.reference_area [m^2]
vehicle.wings.
areas.reference [m^2]
thickness_to_chord [Unitless]
chords.mean_aerodynamic [m]
sweeps.quarter_chord [radians]
taper [Unitless]
flaps.chord [m]
control_surfaces.flap.deflection [radians]
control_surfaces.slat.deflection [radians]
areas.affected [m^2]
control_surfaces.flap.configuration_type [string]
conditions.freestream.
velocity [m/s]
density [kg/m^3]
dynamic_viscosity [N s/m^2]
Outputs:
Cl_max_ls (maximum CL) [Unitless]
Cd_ind (induced drag) [Unitless]
Properties Used:
N/A
"""
# initializing Cl and CDi
Cl_max_ls = 0
Cd_ind = 0
vehicle = geometry
conditions = state.conditions
#unpack
max_lift_coefficient_factor = settings.maximum_lift_coefficient_factor
for wing in vehicle.wings:
if not wing.high_lift: continue
#geometrical data
Sref = vehicle.reference_area
Swing = wing.areas.reference
tc = wing.thickness_to_chord * 100
chord_mac = wing.chords.mean_aerodynamic
sweep = wing.sweeps.quarter_chord
sweep_deg = wing.sweeps.quarter_chord / Units.degree # convert into degrees
taper = wing.taper
# conditions data
V = conditions.freestream.velocity
roc = conditions.freestream.density
nu = conditions.freestream.dynamic_viscosity
#--cl max based on airfoil t_c
Cl_max_ref = -0.0009*tc**3 + 0.0217*tc**2 - 0.0442*tc + 0.7005
#-reynolds number effect
Reyn = V * roc * chord_mac / nu
Re_ref = 9*10**6
op_Clmax = Cl_max_ref * ( Reyn / Re_ref ) **0.1
#wing cl_max to outer panel Cl_max
w_Clmax = op_Clmax* ( 0.919729714285715 -0.044504761904771*taper \
-0.001835900000000*sweep_deg + 0.247071428571446*taper**2 + \
0.003191500000000*taper*sweep_deg -0.000056632142857*sweep_deg**2 \
-0.279166666666676*taper**3 + 0.002300000000000*taper**2*sweep_deg + \
0.000049982142857*taper*sweep_deg**2 -0.000000280000000* sweep_deg**3)
#---FAR stall speed effect---------------
#should be optional based on aircraft being modelled
Cl_max_FAA = 1.1 * w_Clmax
#-----------wing mounted engine ----
Cl_max_w_eng = Cl_max_FAA - 0.2
# Compute CL increment due to Flap
if 'slat' in wing.control_surfaces.keys():
# Compute CL increment due to Slat
slat_angle = wing.control_surfaces.slat.deflection
dcl_slat = compute_slat_lift(slat_angle, sweep)
else:
dcl_slat = 0.
# Compute CL increment due to Flap
if 'flap' in wing.control_surfaces.keys():
flap_type = wing.control_surfaces.flap.configuration_type
flap_chord = wing.control_surfaces.flap.chord_fraction # correct !!!
flap_angle = wing.control_surfaces.flap.deflection
Swf = wing.areas.affected # portion of wing area with flaps
dcl_flap = compute_flap_lift(tc,flap_type,flap_chord,flap_angle,sweep,Sref,Swf)
else:
dcl_flap = 0.0
#results
Cl_max_ls += (Cl_max_w_eng + dcl_slat + dcl_flap) * Swing / Sref
Cd_ind += ( 0.01 ) * Swing / Sref
Cl_max_ls = Cl_max_ls * max_lift_coefficient_factor
return Cl_max_ls, Cd_ind
