RCAIDE.Library.Methods.Aerodynamics.Airfoil_Panel_Method.hess_smith
hess_smith#
- hess_smith(x_coord, y_coord, alpha, Re, npanel)[source]#
Computes the incompressible, inviscid flow over an airfoil of arbitrary shape using the Hess-Smith panel method.
Assumptions: None
- Source: “An introduction to theoretical and computational
aerodynamics”, J. Moran, Wiley, 1984
Inputs x - Vector of x coordinates of the surface [unitess] y - Vector of y coordinates of the surface [unitess] alpha - Airfoil angle of attack [radians] npanel - Number of panels on the airfoil. The number of nodes [unitess]
is equal to npanel+1, and the ith panel goes from node i to node i+1
Outputs cl - Airfoil lift coefficient [unitless] cd - Airfoil drag coefficient [unitless] cm - Airfoil moment coefficient about the c/4 [unitless] x_bar - Vector of x coordinates of the surface nodes [unitless] y_bar - Vector of y coordinates of the surface nodes [unitless] cp - Vector of coefficients of pressure at the nodes [unitless]
Properties Used: N/A