XFOIL         Version 6.94
  
 Calculated polar for: NACA 63-210 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.1708   0.00807   0.00282  -0.0403  0.7899  0.7953
   0.500   0.2220   0.00798   0.00263  -0.0386  0.7506  0.8168
   1.000   0.2742   0.00797   0.00260  -0.0373  0.7217  0.8399
   1.500   0.3249   0.00798   0.00261  -0.0357  0.6871  0.8651
   2.000   0.3730   0.00805   0.00263  -0.0335  0.6387  0.8920
   2.500   0.4174   0.00824   0.00268  -0.0304  0.5643  0.9219
   3.000   0.4504   0.00999   0.00302  -0.0264  0.2334  0.9585
   3.500   0.5146   0.01183   0.00404  -0.0296  0.0391  1.0000
   4.000   0.5696   0.01256   0.00485  -0.0301  0.0370  1.0000
   4.500   0.6233   0.01357   0.00593  -0.0303  0.0364  1.0000
   5.000   0.6742   0.01489   0.00732  -0.0299  0.0366  1.0000
   5.500   0.7261   0.01582   0.00827  -0.0300  0.0281  1.0000
   6.000   0.7735   0.01744   0.00990  -0.0294  0.0235  1.0000
   6.500   0.8244   0.01844   0.01102  -0.0291  0.0216  1.0000
   7.000   0.8752   0.01913   0.01177  -0.0291  0.0168  1.0000
   7.500   0.9210   0.02091   0.01377  -0.0282  0.0147  1.0000
   8.000   0.9724   0.02135   0.01437  -0.0279  0.0135  1.0000
   8.500   1.0204   0.02261   0.01598  -0.0270  0.0056  1.0000
   9.000   1.0690   0.02318   0.01654  -0.0265  0.0040  1.0000
   9.500   1.1079   0.02550   0.01921  -0.0247  0.0037  1.0000
  10.000   1.1377   0.02925   0.02351  -0.0220  0.0035  1.0000
  10.500   1.1522   0.03476   0.02976  -0.0182  0.0035  1.0000
  11.000   1.1365   0.04182   0.03758  -0.0124  0.0035  1.0000
  11.500   1.0951   0.05135   0.04781  -0.0093  0.0036  1.0000
  12.000   1.0501   0.06303   0.05999  -0.0122  0.0037  1.0000
  12.500   1.0012   0.07897   0.07632  -0.0218  0.0037  1.0000