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GOE 451 AIRFOIL (modified line 6) (goe451-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 451 AIRFOIL (modified line 6) (goe451-il)
Reynolds number: 500,000
Max Cl/Cd: 37.28 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe451-il-500000.txt
Download as CSV file: xf-goe451-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 451 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -6.250  -0.2688   0.07114   0.06613  -0.0633   0.0365   0.0080
  -6.000  -0.2600   0.06600   0.06101  -0.0653   0.0345   0.0088
  -5.750  -0.2428   0.06282   0.05784  -0.0681   0.0311   0.0092
  -5.500  -0.2113   0.04484   0.04018  -0.0623   0.0306   0.0110
  -2.500   0.1332   0.01979   0.01261  -0.1035   0.0108   0.0151
  -2.000   0.1898   0.01550   0.00735  -0.1035   0.0106   0.0116
  -1.750   0.2176   0.01384   0.00533  -0.1032   0.0106   0.0111
  -1.500   0.2442   0.01270   0.00402  -0.1029   0.0106   0.0123
  -1.250   0.2699   0.01261   0.00381  -0.1027   0.0106   0.0170
  -1.000   0.2953   0.01271   0.00389  -0.1030   0.0104   0.0216
  -0.750   0.3216   0.01259   0.00376  -0.1033   0.0088   0.0246
   1.500   0.5339   0.01763   0.00826  -0.1030   0.0073   0.0627
   1.750   0.5585   0.01815   0.00895  -0.1034   0.0072   0.1661
   2.000   0.5827   0.01804   0.00894  -0.1034   0.0072   0.1833
   2.250   0.5997   0.01651   0.00916  -0.1018   0.0071   1.0000
   2.500   0.6241   0.01708   0.00970  -0.1018   0.0071   1.0000
   2.750   0.6488   0.01785   0.01053  -0.1018   0.0070   1.0000
   3.000   0.6733   0.01806   0.01075  -0.1017   0.0062   1.0000
   3.250   0.7044   0.02138   0.01225  -0.1032   0.0047   0.0779
   3.500   0.7272   0.02643   0.01723  -0.1035   0.0044   0.0725
   3.750   0.7512   0.02710   0.01807  -0.1030   0.0044   0.0709
   4.000   0.7748   0.02817   0.01931  -0.1026   0.0044   0.0717
   4.250   0.7981   0.02938   0.02072  -0.1021   0.0043   0.0745
   4.500   0.8212   0.03059   0.02218  -0.1016   0.0043   0.0800
   4.750   0.8441   0.03179   0.02363  -0.1010   0.0043   0.0870
   7.750   1.0642   0.05367   0.04946  -0.0912   0.0017   1.0000
   8.000   1.0746   0.05765   0.05367  -0.0899   0.0019   1.0000
   8.250   1.0842   0.06186   0.05808  -0.0888   0.0019   1.0000
   8.500   1.0928   0.06613   0.06253  -0.0876   0.0019   1.0000
   8.750   1.1005   0.07040   0.06698  -0.0865   0.0020   1.0000
   9.000   1.1070   0.07466   0.07141  -0.0854   0.0020   1.0000
   9.250   1.1121   0.07889   0.07581  -0.0843   0.0020   1.0000
   9.500   1.1155   0.08318   0.08026  -0.0832   0.0020   1.0000
   9.750   1.1172   0.08735   0.08459  -0.0822   0.0020   1.0000
  10.000   1.1167   0.09153   0.08892  -0.0811   0.0020   1.0000
  10.250   1.1139   0.09550   0.09303  -0.0800   0.0020   1.0000
  10.500   1.1084   0.09945   0.09710  -0.0789   0.0020   1.0000
  10.750   1.0999   0.10326   0.10103  -0.0779   0.0020   1.0000
  11.000   1.0852   0.10641   0.10427  -0.0759   0.0020   1.0000
  11.250   1.0694   0.10993   0.10787  -0.0752   0.0020   1.0000
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