def compute_max_lift_coeff(vehicle,conditions=None):
""" SUAVE.Methods.Aerodynamics.compute_max_lift_coeff(vehicle):
Computes the maximum lift coefficient associated with an aircraft high lift system
Inputs:
vehicle - SUave type vehicle
conditions - data dictionary with fields:
mach_number - float or 1D array of freestream mach numbers
airspeed - float or 1D array of freestream airspeed
rho - air density
mu - air dynamic_viscosity
geometry - Not used
Outputs:
Cl_max_ls - maximum lift coefficient
Cd_ind - induced drag increment due to high lift device
Assumptions:
if needed
"""
# initializing Cl and CDi
Cl_max_ls = 0
Cd_ind = 0
#unpack
max_lift_coefficient_factor = vehicle.max_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.sweep # convert into degrees
taper = wing.taper
flap_chord = wing.flaps.chord
flap_angle = wing.flaps.angle
slat_angle = wing.slats.angle
Swf = wing.areas.affected #portion of wing area with flaps
flap_type = wing.flaps.type
# conditions data
V = conditions.freestream.velocity
roc = conditions.freestream.density
nu = conditions.freestream.dynamic_viscosity
##Mcr = segment.M
#--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 + 0.247071428571446*taper**2 + \
0.003191500000000*taper*sweep -0.000056632142857*sweep**2 \
-0.279166666666676*taper**3 + 0.002300000000000*taper**2*sweep + \
0.000049982142857*taper*sweep**2 -0.000000280000000* sweep**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 Slat
dcl_slat = compute_slat_lift(slat_angle, sweep)
# Compute CL increment due to Flap
dcl_flap = compute_flap_lift(tc,flap_type,flap_chord,flap_angle,sweep,Sref,Swf)
#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
def compute_max_lift_coeff(vehicle,conditions=None):
"""Computes the maximum lift coefficient associated with an aircraft high lift system
Assumptions:
None
Source:
Unknown
Inputs:
vehicle.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]
flaps.angle [radians]
slats.angle [radians]
areas.affected [m^2]
flaps.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
#unpack
max_lift_coefficient_factor = vehicle.max_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 # convert into degrees
taper = wing.taper
flap_chord = wing.flaps.chord
flap_angle = wing.flaps.angle
slat_angle = wing.slats.angle
Swf = wing.areas.affected #portion of wing area with flaps
flap_type = wing.flaps.type
# 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 + 0.247071428571446*taper**2 + \
0.003191500000000*taper*sweep -0.000056632142857*sweep**2 \
-0.279166666666676*taper**3 + 0.002300000000000*taper**2*sweep + \
0.000049982142857*taper*sweep**2 -0.000000280000000* sweep**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 Slat
dcl_slat = compute_slat_lift(slat_angle, sweep)
# Compute CL increment due to Flap
dcl_flap = compute_flap_lift(tc,flap_type,flap_chord,flap_angle,sweep,Sref,Swf)
#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
def compute_max_lift_coeff(vehicle, conditions=None):
"""Computes the maximum lift coefficient associated with an aircraft high lift system
Assumptions:
None
Source:
Unknown
Inputs:
vehicle.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]
flaps.angle [radians]
slats.angle [radians]
areas.affected [m^2]
flaps.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
#unpack
max_lift_coefficient_factor = vehicle.max_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 # convert into degrees
taper = wing.taper
flap_chord = wing.flaps.chord
flap_angle = wing.flaps.angle
slat_angle = wing.slats.angle
Swf = wing.areas.affected #portion of wing area with flaps
flap_type = wing.flaps.type
# 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 + 0.247071428571446*taper**2 + \
0.003191500000000*taper*sweep -0.000056632142857*sweep**2 \
-0.279166666666676*taper**3 + 0.002300000000000*taper**2*sweep + \
0.000049982142857*taper*sweep**2 -0.000000280000000* sweep**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 Slat
dcl_slat = compute_slat_lift(slat_angle, sweep)
# Compute CL increment due to Flap
dcl_flap = compute_flap_lift(tc, flap_type, flap_chord, flap_angle,
sweep, Sref, Swf)
#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