Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

GOE 286 AIRFOIL (goe286-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 286 AIRFOIL (goe286-il)
Reynolds number: 200,000
Max Cl/Cd: 65.54 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe286-il-200000-n5.txt
Download as CSV file: xf-goe286-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 286 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3983   0.10578   0.10233  -0.0152   1.0000   0.0303
  -9.000  -0.3961   0.10184   0.09842  -0.0186   1.0000   0.0304
  -8.750  -0.3932   0.09788   0.09449  -0.0216   1.0000   0.0305
  -8.500  -0.3904   0.09393   0.09057  -0.0242   1.0000   0.0305
  -8.250  -0.3856   0.08997   0.08665  -0.0239   1.0000   0.0309
  -8.000  -0.3757   0.08744   0.08413  -0.0228   1.0000   0.0315
  -7.750  -0.3697   0.08467   0.08139  -0.0235   1.0000   0.0321
  -7.500  -0.3631   0.08159   0.07834  -0.0256   1.0000   0.0329
  -7.250  -0.3564   0.07860   0.07538  -0.0278   1.0000   0.0345
  -7.000  -0.3492   0.07245   0.06917  -0.0405   0.9983   0.0381
  -6.750  -0.3212   0.06605   0.06262  -0.0490   0.9903   0.0382
  -6.500  -0.2942   0.06023   0.05665  -0.0552   0.9832   0.0383
  -6.250  -0.2802   0.05827   0.05481  -0.0534   0.9763   0.0366
  -6.000  -0.2522   0.05293   0.04931  -0.0593   0.9664   0.0362
  -5.750  -0.2236   0.04757   0.04371  -0.0644   0.9552   0.0361
  -5.500  -0.1977   0.04267   0.03857  -0.0677   0.9420   0.0348
  -5.250  -0.1728   0.03786   0.03344  -0.0699   0.9265   0.0342
  -5.000  -0.1484   0.03413   0.02936  -0.0709   0.9088   0.0353
  -4.750  -0.1240   0.03031   0.02512  -0.0713   0.8902   0.0355
  -4.500  -0.0996   0.02702   0.02140  -0.0711   0.8717   0.0353
  -4.250  -0.0747   0.02433   0.01825  -0.0707   0.8538   0.0354
  -4.000  -0.0493   0.02228   0.01575  -0.0702   0.8363   0.0361
  -3.750  -0.0236   0.02073   0.01378  -0.0695   0.8178   0.0369
  -3.500   0.0024   0.01937   0.01204  -0.0689   0.7980   0.0371
  -3.250   0.0289   0.01825   0.01057  -0.0683   0.7759   0.0373
  -3.000   0.0556   0.01734   0.00935  -0.0678   0.7507   0.0375
  -2.750   0.0824   0.01659   0.00833  -0.0673   0.7222   0.0378
  -2.500   0.1091   0.01599   0.00746  -0.0668   0.6930   0.0380
  -2.250   0.1357   0.01546   0.00668  -0.0663   0.6686   0.0384
  -2.000   0.1623   0.01480   0.00585  -0.0660   0.6483   0.0396
  -1.750   0.1894   0.01444   0.00538  -0.0657   0.6312   0.0408
  -1.500   0.2167   0.01413   0.00495  -0.0655   0.6162   0.0416
  -1.250   0.2443   0.01383   0.00457  -0.0653   0.6030   0.0422
  -1.000   0.2720   0.01358   0.00424  -0.0652   0.5906   0.0430
  -0.750   0.2998   0.01337   0.00396  -0.0651   0.5776   0.0439
  -0.500   0.3278   0.01319   0.00372  -0.0650   0.5637   0.0451
  -0.250   0.3558   0.01306   0.00352  -0.0650   0.5482   0.0464
   0.000   0.3837   0.01295   0.00334  -0.0649   0.5309   0.0483
   0.250   0.4116   0.01286   0.00322  -0.0648   0.5112   0.0527
   0.500   0.4393   0.01281   0.00314  -0.0648   0.4890   0.0641
   0.750   0.4667   0.01275   0.00306  -0.0647   0.4672   0.0982
   1.000   0.4934   0.01240   0.00308  -0.0648   0.4476   0.2499
   1.250   0.5211   0.01071   0.00312  -0.0646   0.4308   1.0000
   1.500   0.5482   0.01092   0.00315  -0.0644   0.4157   1.0000
   1.750   0.5753   0.01114   0.00321  -0.0642   0.4015   1.0000
   2.000   0.6025   0.01135   0.00329  -0.0641   0.3901   1.0000
   2.250   0.6298   0.01153   0.00337  -0.0640   0.3798   1.0000
   2.500   0.6570   0.01174   0.00348  -0.0639   0.3701   1.0000
   2.750   0.6840   0.01196   0.00361  -0.0637   0.3593   1.0000
   3.000   0.7111   0.01217   0.00373  -0.0636   0.3473   1.0000
   3.250   0.7380   0.01240   0.00387  -0.0635   0.3340   1.0000
   3.500   0.7647   0.01264   0.00403  -0.0634   0.3200   1.0000
   3.750   0.7914   0.01289   0.00420  -0.0632   0.3070   1.0000
   4.000   0.8180   0.01314   0.00438  -0.0631   0.2967   1.0000
   4.250   0.8449   0.01335   0.00458  -0.0630   0.2872   1.0000
   4.500   0.8714   0.01361   0.00479  -0.0628   0.2751   1.0000
   4.750   0.8978   0.01388   0.00501  -0.0626   0.2638   1.0000
   5.000   0.9239   0.01417   0.00527  -0.0625   0.2503   1.0000
   5.250   0.9497   0.01449   0.00553  -0.0623   0.2336   1.0000
   5.500   0.9750   0.01489   0.00583  -0.0620   0.2104   1.0000
   5.750   0.9963   0.01581   0.00638  -0.0614   0.1491   1.0000
   6.000   1.0204   0.01635   0.00685  -0.0611   0.1385   1.0000
   6.250   1.0453   0.01675   0.00726  -0.0607   0.1346   1.0000
   6.500   1.0701   0.01714   0.00768  -0.0604   0.1317   1.0000
   6.750   1.0947   0.01754   0.00813  -0.0601   0.1291   1.0000
   7.000   1.1191   0.01796   0.00859  -0.0597   0.1268   1.0000
   7.250   1.1426   0.01844   0.00912  -0.0592   0.1236   1.0000
   7.500   1.1667   0.01884   0.00959  -0.0588   0.1219   1.0000
   7.750   1.1906   0.01924   0.01010  -0.0583   0.1201   1.0000
   8.000   1.2136   0.01971   0.01066  -0.0578   0.1169   1.0000
   8.250   1.2350   0.02033   0.01132  -0.0571   0.1123   1.0000
   8.500   1.2557   0.02098   0.01202  -0.0563   0.1078   1.0000
   8.750   1.2789   0.02137   0.01255  -0.0558   0.1037   1.0000
   9.000   1.3001   0.02192   0.01317  -0.0550   0.0979   1.0000
   9.250   1.3226   0.02233   0.01369  -0.0544   0.0901   1.0000
   9.500   1.3415   0.02303   0.01432  -0.0535   0.0657   1.0000
   9.750   1.3562   0.02407   0.01526  -0.0521   0.0590   1.0000
  10.000   1.3699   0.02510   0.01636  -0.0505   0.0550   1.0000
  10.250   1.3804   0.02630   0.01764  -0.0486   0.0514   1.0000
  10.500   1.3932   0.02719   0.01866  -0.0469   0.0484   1.0000
  10.750   1.4019   0.02825   0.01985  -0.0446   0.0441   1.0000
  11.000   1.4113   0.02935   0.02108  -0.0427   0.0398   1.0000
  11.250   1.4228   0.03034   0.02220  -0.0412   0.0317   1.0000
  11.500   1.4262   0.03204   0.02381  -0.0393   0.0218   1.0000
  11.750   1.4279   0.03400   0.02577  -0.0376   0.0191   1.0000
  12.000   1.4301   0.03608   0.02796  -0.0363   0.0180   1.0000
  12.250   1.4308   0.03846   0.03048  -0.0353   0.0171   1.0000
  12.500   1.4296   0.04119   0.03335  -0.0346   0.0164   1.0000
  12.750   1.4263   0.04435   0.03666  -0.0344   0.0158   1.0000
  13.000   1.4207   0.04801   0.04048  -0.0346   0.0153   1.0000
  13.250   1.4139   0.05205   0.04468  -0.0353   0.0149   1.0000
  13.500   1.4089   0.05596   0.04876  -0.0360   0.0146   1.0000
  13.750   1.4020   0.06019   0.05316  -0.0370   0.0142   1.0000
  14.000   1.3935   0.06471   0.05785  -0.0381   0.0140   1.0000
  14.250   1.3837   0.06951   0.06280  -0.0393   0.0137   1.0000
  14.500   1.3733   0.07452   0.06797  -0.0408   0.0134   1.0000
  14.750   1.3626   0.07974   0.07333  -0.0424   0.0132   1.0000
  15.000   1.3518   0.08510   0.07882  -0.0442   0.0130   1.0000
  15.250   1.3411   0.09053   0.08439  -0.0461   0.0129   1.0000
<< Back to GOE 286 AIRFOIL (goe286-il)

Polar data table (+)

Polar graphs


<< Back to GOE 286 AIRFOIL (goe286-il)