XFOIL         Version 6.94
  
 Calculated polar for: GOE 180 (MVA H.26) 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   0.4870   0.01029   0.00244  -0.0973  0.6350  0.1048
   1.000   0.5911   0.00868   0.00268  -0.0958  0.5774  1.0000
   1.500   0.6450   0.00903   0.00278  -0.0953  0.5543  1.0000
   2.000   0.6991   0.00938   0.00295  -0.0948  0.5351  1.0000
   2.500   0.7531   0.00974   0.00315  -0.0944  0.5177  1.0000
   3.000   0.8072   0.01009   0.00339  -0.0941  0.5031  1.0000
   3.500   0.8590   0.01038   0.00360  -0.0932  0.4716  1.0000
   4.000   0.9074   0.01076   0.00364  -0.0918  0.4020  1.0000
   4.500   0.9571   0.01134   0.00395  -0.0907  0.3337  1.0000
   5.000   0.9995   0.01280   0.00472  -0.0889  0.2225  1.0000
   5.500   1.0449   0.01402   0.00557  -0.0874  0.1488  1.0000
   6.000   1.0820   0.01617   0.00702  -0.0849  0.0396  1.0000
   6.500   1.1249   0.01755   0.00832  -0.0828  0.0060  1.0000
   7.000   1.1713   0.01848   0.00939  -0.0812  0.0059  1.0000
   7.500   1.2155   0.01957   0.01068  -0.0792  0.0062  1.0000
   8.000   1.2559   0.02095   0.01229  -0.0767  0.0066  1.0000
   8.500   1.2944   0.02237   0.01396  -0.0739  0.0072  1.0000
   9.000   1.3239   0.02438   0.01625  -0.0699  0.0080  1.0000
   9.500   1.3353   0.02710   0.01920  -0.0635  0.0086  1.0000
  10.000   1.3467   0.02987   0.02224  -0.0577  0.0097  1.0000
  10.500   1.3418   0.03435   0.02691  -0.0518  0.0105  1.0000
  11.000   1.3482   0.03871   0.03153  -0.0473  0.0119  1.0000
  11.500   1.2214   0.03683   0.03036  -0.0350  0.0112  1.0000
  12.000   1.2648   0.04185   0.03551  -0.0281  0.0152  1.0000