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GOE 55 AIRFOIL (goe55-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 55 AIRFOIL (goe55-il)
Reynolds number: 200,000
Max Cl/Cd: 54.51 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe55-il-200000.txt
Download as CSV file: xf-goe55-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 55 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.6304   0.10791   0.10431   0.0388   1.0000   0.0479
  -8.250  -0.6256   0.10440   0.10083   0.0372   1.0000   0.0491
  -8.000  -0.6219   0.10079   0.09724   0.0349   1.0000   0.0506
  -7.750  -0.5348   0.08487   0.08156   0.0142   1.0000   0.0543
  -7.500  -0.5166   0.08157   0.07825   0.0184   1.0000   0.0556
  -7.250  -0.5109   0.07815   0.07486   0.0184   1.0000   0.0569
  -7.000  -0.5076   0.07429   0.07100   0.0167   1.0000   0.0582
  -6.750  -0.5022   0.06981   0.06653   0.0132   1.0000   0.0599
  -6.500  -0.4935   0.06363   0.06028   0.0030   1.0000   0.0632
  -6.250  -0.4809   0.05447   0.05087  -0.0092   1.0000   0.0649
  -6.000  -0.4699   0.05089   0.04736  -0.0072   1.0000   0.0657
  -5.750  -0.4560   0.04782   0.04433  -0.0064   1.0000   0.0670
  -5.500  -0.4392   0.04452   0.04100  -0.0073   1.0000   0.0696
  -5.000  -0.4160   0.04733   0.04288  -0.0177   1.0000   0.0778
  -4.750  -0.3949   0.04474   0.04034  -0.0179   1.0000   0.0794
  -4.500  -0.3700   0.04226   0.03780  -0.0188   1.0000   0.0825
  -4.250  -0.3338   0.03828   0.03327  -0.0228   1.0000   0.0909
  -4.000  -0.3106   0.03592   0.03102  -0.0230   1.0000   0.0932
  -3.750  -0.2827   0.03397   0.02896  -0.0238   1.0000   0.0990
  -3.500  -0.2511   0.03102   0.02569  -0.0255   1.0000   0.1061
  -3.250  -0.2236   0.02907   0.02372  -0.0260   1.0000   0.1099
  -3.000  -0.1798   0.02292   0.01636  -0.0276   1.0000   0.0781
  -2.750  -0.1493   0.02087   0.01412  -0.0278   1.0000   0.0735
  -2.500  -0.1185   0.01861   0.01157  -0.0280   1.0000   0.0700
  -2.250  -0.0878   0.01682   0.00949  -0.0280   1.0000   0.0683
  -2.000  -0.0583   0.01566   0.00820  -0.0279   1.0000   0.0686
  -1.750  -0.0292   0.01481   0.00732  -0.0278   1.0000   0.0702
  -1.500  -0.0002   0.01416   0.00667  -0.0277   1.0000   0.0731
  -1.250   0.0290   0.01347   0.00599  -0.0275   1.0000   0.0751
  -1.000   0.0536   0.01404   0.00542  -0.0252   0.5982   0.0772
  -0.750   0.0805   0.01398   0.00494  -0.0248   0.5106   0.0812
  -0.500   0.1089   0.01392   0.00468  -0.0247   0.4808   0.0876
  -0.250   0.1374   0.01367   0.00438  -0.0247   0.4635   0.0972
   0.000   0.1665   0.01334   0.00403  -0.0247   0.4503   0.1140
   0.250   0.1959   0.01284   0.00375  -0.0250   0.4396   0.1686
   0.500   0.2152   0.01040   0.00360  -0.0229   0.4319   1.0000
   0.750   0.2440   0.01060   0.00360  -0.0227   0.4215   1.0000
   1.000   0.2725   0.01083   0.00364  -0.0226   0.4098   1.0000
   1.500   0.3296   0.01125   0.00382  -0.0224   0.3904   1.0000
   1.750   0.3580   0.01154   0.00396  -0.0223   0.3834   1.0000
   2.000   0.3867   0.01171   0.00412  -0.0222   0.3754   1.0000
   2.250   0.4149   0.01200   0.00428  -0.0220   0.3671   1.0000
   2.500   0.4434   0.01211   0.00440  -0.0219   0.3564   1.0000
   2.750   0.4717   0.01228   0.00452  -0.0218   0.3449   1.0000
   3.000   0.4998   0.01249   0.00465  -0.0216   0.3343   1.0000
   3.250   0.5282   0.01260   0.00478  -0.0215   0.3235   1.0000
   3.500   0.5564   0.01270   0.00488  -0.0214   0.3106   1.0000
   3.750   0.5846   0.01274   0.00494  -0.0212   0.2957   1.0000
   4.000   0.6129   0.01281   0.00503  -0.0211   0.2817   1.0000
   4.250   0.6412   0.01285   0.00511  -0.0209   0.2656   1.0000
   4.500   0.6695   0.01288   0.00516  -0.0208   0.2451   1.0000
   4.750   0.6978   0.01297   0.00522  -0.0207   0.2158   1.0000
   5.000   0.7255   0.01331   0.00538  -0.0207   0.1766   1.0000
   5.250   0.7527   0.01406   0.00588  -0.0207   0.1314   1.0000
   5.500   0.7798   0.01506   0.00663  -0.0208   0.0878   1.0000
   5.750   0.8072   0.01583   0.00732  -0.0207   0.0743   1.0000
   6.000   0.8346   0.01650   0.00802  -0.0205   0.0686   1.0000
   6.250   0.8610   0.01746   0.00895  -0.0204   0.0646   1.0000
   6.500   0.8878   0.01820   0.00977  -0.0201   0.0622   1.0000
   6.750   0.9143   0.01900   0.01063  -0.0199   0.0597   1.0000
   7.000   0.9402   0.01993   0.01160  -0.0196   0.0579   1.0000
   7.250   0.9656   0.02104   0.01274  -0.0192   0.0564   1.0000
   7.500   0.9902   0.02254   0.01425  -0.0188   0.0552   1.0000
   7.750   1.0148   0.02420   0.01601  -0.0184   0.0542   1.0000
   8.000   1.0400   0.02541   0.01742  -0.0180   0.0533   1.0000
   8.250   1.0647   0.02681   0.01903  -0.0176   0.0522   1.0000
   8.500   1.0885   0.02856   0.02101  -0.0171   0.0515   1.0000
   8.750   1.1113   0.03056   0.02328  -0.0166   0.0509   1.0000
   9.000   1.1330   0.03270   0.02572  -0.0162   0.0501   1.0000
   9.250   1.1537   0.03481   0.02807  -0.0158   0.0490   1.0000
   9.500   1.1730   0.03715   0.03065  -0.0153   0.0481   1.0000
   9.750   1.1889   0.04028   0.03414  -0.0149   0.0478   1.0000
  10.000   1.2007   0.04409   0.03837  -0.0144   0.0477   1.0000
  10.250   1.2082   0.04822   0.04290  -0.0141   0.0474   1.0000
  10.500   1.2106   0.05269   0.04776  -0.0140   0.0472   1.0000
  10.750   1.1240   0.06939   0.06561  -0.0216   0.0510   1.0000
  11.000   1.0613   0.08894   0.08532  -0.0413   0.0532   1.0000
  11.250   1.0402   0.09913   0.09554  -0.0470   0.0554   1.0000
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