Source code for RCAIDE.Library.Methods.Aerodynamics.Airfoil_Panel_Method.chordwise_distribution
# RCAIDE/Methods/Aerodynamics/Airfoil_Panel_Method/chordwise_distribution
#
#
# Created: May 2024, Niranjan Nanjappa
# ----------------------------------------------------------------------------------------------------------------------
# IMPORT
# ----------------------------------------------------------------------------------------------------------------------
# RCAIDE imports
from RCAIDE.Framework.Core import Data
# package imports
import numpy as np
# ----------------------------------------------------------------------------------------------------------------------
# chordwise_distribution
# ----------------------------------------------------------------------------------------------------------------------
[docs]
def chordwise_distribution(x,y,cp,al,npanel,cf,vt):
"""Compute chordwise distribution of lift and drag for
an airfoil
Assumptions:
None
Source:
Anderson, J. D. "Fundamentals of aerodynamics 6th ed". MCGRAW-HILL EDUCATION (2017)
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
cf - skin friction coeff over the surface of airfoil
vt - tangential velocity at all nodes
Outputs:
fD - chordwise distribution of drag
fL - chordwise distribution of lift
Properties Used:
N/A
"""
al = al.T
dx = x[1:]-x[:-1]
dy = y[1:]-y[:-1]
LE_ind = int(npanel/2)
dx_l = -np.flip(dx[0:LE_ind], axis=0)
dx_u = dx[LE_ind:]
dy_l = -np.flip(dy[0:LE_ind], axis=0)
dy_u = dy[LE_ind:]
cp_l = np.flip(cp[0:LE_ind], axis=0)
cp_u = cp[LE_ind:]
cf = cf*(vt/np.abs(vt))
cf_l = -np.flip(cf[0:LE_ind], axis=0)
cf_u = cf[LE_ind:]
dcn = cp_l*dx_l - cp_u*dx_u + cf_u*dy_u - cf_l*dy_l
dca = cp_u*dy_u - cp_l*dy_l + cf_u*dx_u + cf_l*dx_l
dcl = dcn*np.cos(al) - dca*np.sin(al)
dcd = dcn*np.sin(al) + dca*np.cos(al)
cl = np.sum(dcl, axis=0)
cd = np.sum(dcd, axis=0)
fL = dcl/cl
fD = dcd/cd
return fL,fD
