XFOIL         Version 6.94
  
 Calculated polar for: GOE 242 (MVA PR.2) 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
 ------- -------- --------- --------- -------- ------- -------
   0.500   1.0849   0.01291   0.00611  -0.2222  0.5478  0.4842
   1.000   1.1422   0.01311   0.00627  -0.2227  0.5415  0.4904
   1.500   1.1995   0.01338   0.00654  -0.2232  0.5358  0.4957
   2.000   1.2564   0.01352   0.00669  -0.2236  0.5303  0.5020
   2.500   1.3134   0.01369   0.00683  -0.2241  0.5246  0.5069
   3.500   1.4134   0.01371   0.00665  -0.2223  0.4611  0.5183
   4.000   1.4501   0.01486   0.00724  -0.2194  0.3695  0.5232
   4.500   1.4742   0.01704   0.00881  -0.2147  0.2825  0.5273
   5.000   1.4887   0.01931   0.01062  -0.2084  0.2097  0.5321
   6.000   1.4572   0.02762   0.01809  -0.1894  0.0044  0.5391
   6.500   1.4848   0.02963   0.02018  -0.1868  0.0038  0.5444
   7.500   1.5324   0.03465   0.02548  -0.1814  0.0041  0.5544
   8.000   1.5522   0.03769   0.02869  -0.1788  0.0044  0.5599
   8.500   1.5679   0.04124   0.03241  -0.1761  0.0049  0.5662
   9.000   1.5789   0.04544   0.03680  -0.1734  0.0052  0.5714
   9.500   1.5916   0.04966   0.04122  -0.1712  0.0058  0.5773
  10.000   1.5987   0.05476   0.04655  -0.1691  0.0064  0.5832
  10.500   1.6053   0.06008   0.05206  -0.1672  0.0071  0.5895
  11.000   1.6107   0.06574   0.05795  -0.1656  0.0080  0.5958
  11.500   1.6136   0.07179   0.06421  -0.1641  0.0089  0.6031
  12.000   1.6124   0.07849   0.07113  -0.1628  0.0101  0.6106
  12.500   1.6159   0.08473   0.07761  -0.1618  0.0114  0.6185
  13.000   1.6004   0.09362   0.08674  -0.1609  0.0124  0.6255