XFOIL         Version 6.94
  
 Calculated polar for: GOE 376 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
 ------- -------- --------- --------- -------- ------- -------
   1.500   0.6308   0.00324  -0.00438  -0.0823  0.5928  0.0563
   2.000   0.6784   0.00299  -0.00476  -0.0804  0.5533  0.0521
   2.500   0.7255   0.00290  -0.00498  -0.0786  0.5193  0.0521
   3.000   0.7734   0.00289  -0.00508  -0.0772  0.4899  0.0575
   4.000   0.8922   0.00248  -0.00440  -0.0801  0.3881  1.0000
   4.500   0.9299   0.00289  -0.00452  -0.0771  0.2755  1.0000
   5.000   0.9684   0.00347  -0.00441  -0.0744  0.2004  1.0000
   5.500   1.0061   0.00420  -0.00410  -0.0716  0.1269  1.0000
   6.000   1.0322   0.00582  -0.00311  -0.0670  0.0077  1.0000
   6.500   1.0741   0.00643  -0.00235  -0.0646  0.0078  1.0000
   7.000   1.1142   0.00714  -0.00143  -0.0619  0.0087  1.0000
   7.500   1.1499   0.00815  -0.00017  -0.0584  0.0097  1.0000
   8.000   1.1818   0.00935   0.00132  -0.0543  0.0113  1.0000
   8.500   1.2041   0.01087   0.00311  -0.0488  0.0131  1.0000
   9.000   1.2080   0.01281   0.00526  -0.0409  0.0147  1.0000
   9.500   1.1994   0.01575   0.00841  -0.0328  0.0166  1.0000
  10.000   1.2019   0.01914   0.01200  -0.0271  0.0189  1.0000