Source code for RCAIDE.Library.Methods.Aerodynamics.AERODAS.post_stall_coefficients
# RCAIDE/Library/Methods/Aerdoynamics/AERODAS/post_stall_coefficients.py
#
#
# Created: Jul 2024, M. Clarke
# ----------------------------------------------------------------------------------------------------------------------
# Imports
# ----------------------------------------------------------------------------------------------------------------------
from RCAIDE.Framework.Core import Data, Units
# python imports
import numpy as np
# ----------------------------------------------------------------------------------------------------------------------
# Post Stall Coefficients
# ----------------------------------------------------------------------------------------------------------------------
[docs]
def post_stall_coefficients(state,settings,geometry):
"""Uses the AERODAS method to determine poststall parameters for lift and drag for a single wing
Assumptions:
None
Source:
NASA TR: "Models of Lift and Drag Coefficients of Stalled and Unstalled Airfoils in
Wind Turbines and Wind Tunnels" by D. A. Spera
Inputs:
settings.section_zero_lift_angle_of_attack [radians]
geometry.
aspect_ratio [Unitless]
thickness_to_chord [Unitless]
section.angle_attack_max_prestall_lift [radians]
pre_stall_maximum_lift_drag_coefficient [Unitless]
pre_stall_maximum_drag_coefficient_angle [Unitless]
state.conditions.aerodynamics.angle_of_attack [radians]
Outputs:
CL2 (coefficient of lift) [Unitless]
CD2 (coefficient of drag) [Unitless]
(packed in state.conditions.aerodynamics.post_stall_coefficients[geometry.tag])
Properties Used:
N/A
"""
# unpack inputs
wing = geometry
A0 = settings.section_zero_lift_angle_of_attack
AR = wing.aspect_ratio
t_c = wing.thickness_to_chord
ACL1 = wing.section.angle_attack_max_prestall_lift
CD1max = wing.pre_stall_maximum_lift_drag_coefficient
ACD1 = wing.pre_stall_maximum_drag_coefficient_angle
alpha = state.conditions.aerodynamics.angle_of_attack
if wing.vertical == True:
alpha = 0. * np.ones_like(alpha)
# Equation 9a and b
F1 = 1.190*(1.0-(t_c*t_c))
F2 = 0.65 + 0.35*np.exp(-(9.0/AR)**2.3)
# Equation 10b and c
G1 = 2.3*np.exp(-(0.65*t_c)**0.9)
G2 = 0.52 + 0.48*np.exp(-(6.5/AR)**1.1)
# Equation 8a and b
CL2max = F1*F2
CD2max = G1*G2
# Equation 11d
RCL2 = 1.632-CL2max
# Equation 11e
N2 = 1 + CL2max/RCL2
# Equation 11a,b,c
con1 = np.logical_and(0<alpha,alpha<ACL1)
con2 = np.logical_and(ACL1<=alpha,alpha<=(92.0*Units.deg))
con3 = alpha>=(92.0*Units.deg)
CL2 = np.zeros_like(alpha)
CL2[con1] = 0
CL2[con2] = -0.032*(alpha[con2]/Units.deg-92.0) - RCL2*((92.*Units.deg-alpha[con2])/(51.0*Units.deg))**N2
CL2[con3] = -0.032*(alpha[con3]/Units.deg-92.0) + RCL2*((alpha[con3]-92.*Units.deg)/(51.0*Units.deg))**N2
# If alpha is negative flip things for lift
alphan = - alpha+2*A0
con1 = np.logical_and(0<alphan, alphan<ACL1)
con2 = np.logical_and(ACL1<=alphan, alphan<=(92.0*Units.deg))
con3 = alphan>=(92.0*Units.deg)
CL2[con1] = 0.
CL2[con2] = 0.032*(alphan[con2]/Units.deg-92.0) + RCL2*((92.*Units.deg-alphan[con2])/(51.0*Units.deg))**N2
CL2[con3] = 0.032*(alphan[con3]/Units.deg-92.0) - RCL2*((alphan[con3]-92.*Units.deg)/(51.0*Units.deg))**N2
# Equation 12a
con1 = np.logical_and((2*A0-ACL1)<alpha, alpha<ACL1)
con2 = alpha>ACD1
CD2 = 0.0 * np.ones_like(alpha)
CD2[con1] = 0.
CD2[con2] = CD1max[con2] + (CD2max - CD1max[con2]) * np.sin((alpha[con2]-ACD1[con2])/(np.pi/2-ACD1[con2]))
# If alpha is negative flip things for drag
alphan = -alpha + 2*A0
con1 = np.logical_and((2*A0-ACL1)<alphan,alphan<ACL1)
con2 = alphan>=ACD1
CD2[con1] = 0.
CD2[con2] = CD1max[con2] + (CD2max - CD1max[con2]) * np.sin((alphan[con2]-ACD1[con2])/(np.pi/2-ACD1[con2]))
# Pack outputs
wing_result = Data(
lift_coefficient = CL2,
drag_coefficient = CD2
)
state.conditions.aerodynamics.post_stall_coefficients[wing.tag] = wing_result
return CL2, CD2
