XFOIL         Version 6.94
  
 Calculated polar for: EPPLER 407 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.5252   0.01134   0.00658  -0.1452  0.8671  0.6533
   0.500   0.5757   0.01122   0.00648  -0.1441  0.8576  0.6583
   1.500   0.7298   0.01017   0.00545  -0.1529  0.8448  0.6685
   2.000   0.7745   0.00998   0.00537  -0.1505  0.8303  0.6730
   2.500   0.8341   0.00970   0.00515  -0.1511  0.8161  0.6782
   3.000   0.8927   0.00948   0.00496  -0.1516  0.7980  0.6834
   3.500   0.9403   0.00937   0.00488  -0.1497  0.7706  0.6882
   4.000   0.9878   0.00937   0.00487  -0.1478  0.7332  0.6931
   4.500   1.0216   0.00967   0.00500  -0.1430  0.6726  0.6983
   5.000   1.0409   0.01044   0.00543  -0.1355  0.5938  0.7037
   5.500   1.0564   0.01154   0.00616  -0.1278  0.5101  0.7087
   6.000   1.0715   0.01279   0.00709  -0.1203  0.4270  0.7142
   6.500   1.0911   0.01421   0.00815  -0.1142  0.3435  0.7202
   7.000   1.1157   0.01570   0.00928  -0.1093  0.2643  0.7258
   7.500   1.1424   0.01720   0.01048  -0.1051  0.1931  0.7312
   8.000   1.1699   0.01879   0.01182  -0.1012  0.1354  0.7372
   8.500   1.1985   0.02044   0.01329  -0.0975  0.0910  0.7441
   9.000   1.2258   0.02225   0.01495  -0.0939  0.0579  0.7502
   9.500   1.2510   0.02422   0.01689  -0.0901  0.0385  0.7567
  10.000   1.2761   0.02630   0.01902  -0.0864  0.0281  0.7639
  10.500   1.2993   0.02853   0.02133  -0.0828  0.0216  0.7714
  11.000   1.3235   0.03071   0.02371  -0.0795  0.0169  0.7800
  13.000   1.3742   0.04469   0.03860  -0.0657  0.0073  0.8204
  13.500   1.3795   0.04941   0.04362  -0.0631  0.0065  0.8354
  14.000   1.3744   0.05563   0.05011  -0.0609  0.0060  0.8540
  14.500   1.3680   0.06225   0.05716  -0.0591  0.0057  0.8865