Source code for RCAIDE.Library.Methods.Aerodynamics.Airfoil_Panel_Method.aero_coeff
# RCAIDE/Methods/Aerodynamics/Airfoil_Panel_Method/aero_coeff.py
#
#
# Created: Dec 2023, M. Clarke
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# IMPORT
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# RCAIDE imports
from RCAIDE.Framework.Core import Data
# pacakge imports
import numpy as np
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# aero_coeff
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[docs]
def aero_coeff(x,y,cp,al,npanel):
"""Compute airfoil force and moment coefficients about
the quarter chord point
Assumptions:
None
Source:
None
Inputs:
x - Vector of x coordinates of the surface nodes
y - Vector of y coordinates of the surface nodes
cp - Vector of coefficients of pressure at the nodes
al - Angle of attack in radians
npanel - Number of panels on the airfoil
Outputs:
cl - Airfoil lift coefficient
cd - Airfoil drag coefficient
cm - Airfoil moment coefficient about the c/4
Properties Used:
N/A
"""
dx = x[1:]-x[:-1]
dy = y[1:]-y[:-1]
xa = 0.5*(x[1:] +x[:-1])-0.25
ya = 0.5*(y[1:] +y[:-1])
dcn = -cp*dx
dca = cp*dy
# compute differential forces
cn = np.sum(dcn,axis=0).T
ca = np.sum(dca,axis=0).T
cm = np.sum((-dcn*xa + dca*ya),axis=0).T
# orient normal and axial forces
cl = cn*np.cos(al) - ca*np.sin(al)
cdpi = cn*np.sin(al) + ca*np.cos(al)
# pack results
AERO_RES = Data(
cl = cl,
cdpi = cdpi,
cm = cm)
return AERO_RES
