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GOE 9K AIRFOIL (goe09k-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 9K AIRFOIL (goe09k-il)
Reynolds number: 200,000
Max Cl/Cd: 29.31 at α=1.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe09k-il-200000.txt
Download as CSV file: xf-goe09k-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 9K 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5855   0.09677   0.09323   0.0055   1.0000   0.0167
  -7.750  -0.5833   0.09339   0.08990   0.0040   1.0000   0.0173
  -7.500  -0.5819   0.09019   0.08674   0.0025   1.0000   0.0178
  -7.250  -0.5775   0.08672   0.08329  -0.0002   1.0000   0.0183
  -7.000  -0.5694   0.08298   0.07957  -0.0040   1.0000   0.0188
  -6.750  -0.5567   0.07909   0.07568  -0.0090   1.0000   0.0195
  -6.500  -0.5381   0.07537   0.07189  -0.0149   1.0000   0.0200
  -6.250  -0.5176   0.07174   0.06815  -0.0195   1.0000   0.0204
  -6.000  -0.4994   0.06797   0.06426  -0.0222   1.0000   0.0205
  -5.750  -0.4988   0.06085   0.05720  -0.0229   1.0000   0.0217
  -5.500  -0.4889   0.05699   0.05332  -0.0230   1.0000   0.0232
  -5.250  -0.4719   0.05307   0.04928  -0.0245   1.0000   0.0252
  -5.000  -0.4516   0.04908   0.04513  -0.0262   1.0000   0.0275
  -4.750  -0.4253   0.04522   0.04099  -0.0279   1.0000   0.0308
  -4.500  -0.3987   0.04141   0.03675  -0.0289   1.0000   0.0321
  -4.250  -0.3844   0.03771   0.03310  -0.0288   1.0000   0.0358
  -4.000  -0.3599   0.03432   0.02935  -0.0288   1.0000   0.0408
  -3.750  -0.3354   0.03182   0.02640  -0.0284   1.0000   0.0487
  -3.500  -0.3128   0.02936   0.02359  -0.0279   1.0000   0.0589
  -3.250  -0.2930   0.02618   0.02044  -0.0276   1.0000   0.0708
  -3.000  -0.2707   0.02395   0.01801  -0.0270   1.0000   0.0911
  -2.250  -0.1827   0.01589   0.00834  -0.0217   1.0000   0.0299
  -2.000  -0.1562   0.01378   0.00595  -0.0203   1.0000   0.0242
  -1.750  -0.1309   0.01230   0.00434  -0.0191   1.0000   0.0227
  -1.500  -0.1065   0.01114   0.00315  -0.0181   1.0000   0.0277
  -1.250  -0.0614   0.00688   0.00190  -0.0215   1.0000   1.0000
  -1.000  -0.0374   0.00689   0.00151  -0.0206   1.0000   1.0000
  -0.750  -0.0135   0.00690   0.00130  -0.0198   1.0000   1.0000
  -0.500   0.0103   0.00692   0.00118  -0.0190   1.0000   1.0000
  -0.250   0.0341   0.00695   0.00111  -0.0183   1.0000   1.0000
   0.000   0.0579   0.00698   0.00106  -0.0175   1.0000   1.0000
   0.250   0.0815   0.00703   0.00108  -0.0168   1.0000   1.0000
   0.500   0.1051   0.00709   0.00114  -0.0160   1.0000   1.0000
   0.750   0.1285   0.00716   0.00125  -0.0152   1.0000   1.0000
   1.000   0.1519   0.00724   0.00139  -0.0145   1.0000   1.0000
   1.250   0.1751   0.00733   0.00157  -0.0137   1.0000   1.0000
   1.500   0.1983   0.00743   0.00182  -0.0129   1.0000   1.0000
   1.750   0.2213   0.00755   0.00219  -0.0121   1.0000   1.0000
   2.000   0.3323   0.01188   0.00350  -0.0283   0.0240   1.0000
   2.250   0.3543   0.01348   0.00513  -0.0267   0.0219   1.0000
   2.500   0.3798   0.01486   0.00665  -0.0253   0.0239   1.0000
   2.750   0.4071   0.01711   0.00909  -0.0239   0.0303   1.0000
   3.750   0.5192   0.02831   0.02187  -0.0171   0.0778   1.0000
   4.000   0.5407   0.03032   0.02430  -0.0159   0.0608   1.0000
   4.250   0.5615   0.03314   0.02744  -0.0149   0.0505   1.0000
   4.500   0.5824   0.03611   0.03078  -0.0138   0.0426   1.0000
   4.750   0.6020   0.03959   0.03457  -0.0129   0.0373   1.0000
   5.000   0.6178   0.04334   0.03836  -0.0125   0.0328   1.0000
   5.250   0.6390   0.04713   0.04264  -0.0117   0.0308   1.0000
   5.500   0.6591   0.05105   0.04689  -0.0116   0.0274   1.0000
   5.750   0.6750   0.05514   0.05118  -0.0119   0.0248   1.0000
   6.000   0.6860   0.05925   0.05537  -0.0119   0.0227   1.0000
   7.250   0.7352   0.08305   0.07971  -0.0218   0.0194   1.0000
   7.500   0.7404   0.08742   0.08411  -0.0252   0.0190   1.0000
   7.750   0.7417   0.09163   0.08832  -0.0285   0.0187   1.0000
   8.000   0.7395   0.09572   0.09237  -0.0316   0.0184   1.0000
   8.250   0.7393   0.09975   0.09637  -0.0341   0.0180   1.0000
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