XFOIL         Version 6.94
  
 Calculated polar for: GOE 147 (MVA H.6) 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   0.3447   0.00960   0.00219  -0.0572  0.5552  0.1127
   0.500   0.4016   0.00957   0.00201  -0.0569  0.5231  0.1430
   1.000   0.4580   0.00962   0.00195  -0.0568  0.4857  0.1479
   1.500   0.5143   0.00971   0.00194  -0.0567  0.4355  0.1568
   2.000   0.5697   0.01001   0.00206  -0.0567  0.3661  0.1858
   2.500   0.6253   0.01017   0.00230  -0.0568  0.3266  0.2793
   3.000   0.6740   0.00893   0.00255  -0.0554  0.3030  1.0000
   3.500   0.7290   0.00945   0.00276  -0.0554  0.2586  1.0000
   4.000   0.7834   0.01009   0.00303  -0.0553  0.2030  1.0000
   4.500   0.8352   0.01142   0.00385  -0.0551  0.1085  1.0000
   5.000   0.8858   0.01302   0.00500  -0.0546  0.0068  1.0000
   5.500   0.9403   0.01361   0.00566  -0.0542  0.0066  1.0000
   6.000   0.9946   0.01426   0.00645  -0.0536  0.0079  1.0000
   6.500   1.0480   0.01513   0.00751  -0.0529  0.0087  1.0000
   7.000   1.0985   0.01663   0.00930  -0.0518  0.0086  1.0000
   7.500   1.1442   0.01885   0.01179  -0.0502  0.0083  1.0000
   8.000   1.1840   0.02165   0.01481  -0.0481  0.0081  1.0000
   8.500   1.2178   0.02491   0.01824  -0.0454  0.0078  1.0000
   9.000   1.2474   0.02886   0.02226  -0.0422  0.0074  1.0000
   9.500   1.2757   0.03280   0.02661  -0.0390  0.0064  1.0000
  10.000   1.2233   0.02883   0.02308  -0.0279  0.0053  1.0000
  10.500   1.1982   0.03232   0.02711  -0.0224  0.0046  1.0000