XFOIL         Version 6.94
  
 Calculated polar for: GOE 405 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   0.6902   0.00976   0.00292  -0.1415  0.6765  0.1829
   0.500   0.7410   0.00985   0.00305  -0.1405  0.6623  0.2247
   1.000   0.7906   0.00982   0.00304  -0.1392  0.6475  0.2437
   1.500   0.8395   0.00984   0.00303  -0.1379  0.6313  0.2671
   2.000   0.8859   0.00980   0.00305  -0.1361  0.6123  0.3066
   2.500   0.9626   0.00871   0.00333  -0.1414  0.5885  1.0000
   3.000   1.0040   0.00899   0.00346  -0.1385  0.5649  1.0000
   3.500   1.0304   0.00951   0.00365  -0.1325  0.5088  1.0000
   4.000   1.0362   0.01059   0.00419  -0.1226  0.4231  1.0000
   4.500   1.0627   0.01130   0.00468  -0.1171  0.3845  1.0000
   5.000   1.0827   0.01233   0.00536  -0.1107  0.3251  1.0000
   5.500   1.0932   0.01396   0.00646  -0.1031  0.2383  1.0000
   6.000   1.1117   0.01546   0.00762  -0.0973  0.1821  1.0000
   6.500   1.0916   0.01916   0.01053  -0.0863  0.0054  1.0000
   7.000   1.1240   0.02023   0.01167  -0.0832  0.0053  1.0000
   7.500   1.1555   0.02141   0.01296  -0.0802  0.0059  1.0000
   8.000   1.1848   0.02279   0.01445  -0.0770  0.0067  1.0000
   8.500   1.2120   0.02439   0.01619  -0.0738  0.0075  1.0000
   9.000   1.2347   0.02640   0.01838  -0.0703  0.0087  1.0000
   9.500   1.2549   0.02873   0.02088  -0.0668  0.0097  1.0000
  10.000   1.2634   0.03213   0.02449  -0.0627  0.0105  1.0000
  10.500   1.2609   0.03677   0.02933  -0.0587  0.0110  1.0000
  11.000   1.2656   0.04114   0.03392  -0.0557  0.0120  1.0000
  11.500   1.2535   0.04744   0.04044  -0.0526  0.0130  1.0000
  12.000   1.2490   0.05331   0.04653  -0.0500  0.0147  1.0000
  12.500   1.2508   0.05804   0.05138  -0.0461  0.0183  1.0000