XFOIL         Version 6.94
  
 Calculated polar for: GOE 137 (MVA H.15) 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   0.4493   0.00953   0.00208  -0.0833  0.6303  0.0477
   0.500   0.5050   0.00872   0.00201  -0.0837  0.6118  0.3080
   2.500   0.7252   0.00783   0.00221  -0.0831  0.5243  1.0000
   3.000   0.7791   0.00817   0.00232  -0.0829  0.4839  1.0000
   3.500   0.8281   0.00908   0.00258  -0.0820  0.3742  1.0000
   4.000   0.8717   0.01094   0.00347  -0.0808  0.1883  1.0000
   4.500   0.9193   0.01224   0.00432  -0.0799  0.1048  1.0000
   5.000   0.9642   0.01389   0.00554  -0.0784  0.0066  1.0000
   5.500   1.0159   0.01451   0.00626  -0.0777  0.0076  1.0000
   6.000   1.0669   0.01522   0.00710  -0.0767  0.0094  1.0000
   6.500   1.1154   0.01621   0.00830  -0.0753  0.0104  1.0000
   7.000   1.1590   0.01772   0.01001  -0.0731  0.0111  1.0000
   7.500   1.1923   0.02008   0.01254  -0.0696  0.0118  1.0000
   8.000   1.2246   0.02222   0.01484  -0.0658  0.0132  1.0000
   8.500   1.2483   0.02499   0.01773  -0.0607  0.0154  1.0000
   9.000   1.2785   0.02868   0.02142  -0.0565  0.0168  1.0000
   9.500   1.3266   0.03295   0.02568  -0.0555  0.0143  1.0000
  10.000   1.3360   0.03572   0.02891  -0.0497  0.0109  1.0000
  10.500   1.3375   0.04011   0.03363  -0.0446  0.0070  1.0000
  11.000   1.3293   0.04488   0.03882  -0.0398  0.0050  1.0000
  11.500   1.3218   0.05138   0.04568  -0.0365  0.0039  1.0000
  12.000   1.3020   0.05763   0.05244  -0.0350  0.0034  1.0000
  12.500   1.2851   0.06541   0.06058  -0.0351  0.0031  1.0000
  13.000   1.2674   0.07409   0.06950  -0.0360  0.0029  1.0000
  14.000   1.2218   0.09508   0.09101  -0.0406  0.0027  1.0000
  14.500   1.1941   0.10706   0.10331  -0.0471  0.0027  1.0000