Question

The purpose of this work is to compute the drag polar of two airfoils under viscous, incompressible flow conditions. You are asked to analyze drag polar for the following airfoils; NACA 23012 airfoil section, Reynolds number based on chord = 3 Million. NACA 2412 airfoil section, Reynolds number based on chord = 3 Million. The angle of attack should cover -8 degrees to 10 degrees This problem may be solved using XFOIL. Plot Cl, Cm, and Cd vs α of both airfoils, on the same plot. Compare the analysis results with airfoil test data found in the National Advisory Committee for Aeronautics, Report No.824; http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19930090976.pdf, and provide comments. Please overlay your analysis results (Cl, Cm, and Cd) on the test data plots.

Answer 1

The following code in XFOIL gets us the drag polar, Cl-alpha curve and Cm alpha curve

XFOIL c> NACA 23012
XFOIL c> pane
XFOIL c> OPER
.OPERi c> Re
Enter new Reynolds Number r> 1000000
.OPERi c> ASeq
Enter first alfa value (deg) r> -8
Enter last alfa value (deg) r> 10
Enter alfa increment (deg) r> 3

the following is the plot

by changing NACA value we get polot for NACA 2412
the plot is as under

From the given literature
we can see that the graphs obtained by the XFOIL software are based on mathematical equations which have already assumned many simplifications hence the graphs arent ireregular like they are in the experiments