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GOE 622 AIRFOIL (goe622-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 622 AIRFOIL (goe622-il)
Reynolds number: 50,000
Max Cl/Cd: 35.5 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe622-il-50000.txt
Download as CSV file: xf-goe622-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 622 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4753   0.09690   0.09023  -0.0042   1.0000   0.2433
  -7.500  -0.4738   0.09359   0.08700  -0.0044   1.0000   0.2540
  -7.250  -0.4786   0.09082   0.08434  -0.0051   1.0000   0.2656
  -6.750  -0.4681   0.08405   0.07765  -0.0042   1.0000   0.2962
  -6.500  -0.4627   0.08088   0.07454  -0.0034   1.0000   0.3155
  -6.250  -0.4616   0.07783   0.07157  -0.0029   1.0000   0.3378
  -6.000  -0.4512   0.07471   0.06849   0.0006   1.0000   0.3699
  -5.750  -0.4551   0.07268   0.06656   0.0025   1.0000   0.4065
  -5.500  -0.4365   0.06888   0.06276   0.0077   1.0000   0.4420
  -5.250  -0.4287   0.06601   0.05995   0.0120   1.0000   0.4836
  -5.000  -0.4175   0.05414   0.04756  -0.0273   1.0000   0.2655
  -4.750  -0.3729   0.04152   0.03325  -0.0405   1.0000   0.1458
  -4.500  -0.3515   0.03762   0.02892  -0.0405   1.0000   0.1428
  -4.250  -0.3282   0.03446   0.02500  -0.0404   1.0000   0.1478
  -4.000  -0.3063   0.03176   0.02215  -0.0397   1.0000   0.1549
  -3.750  -0.2814   0.02925   0.01909  -0.0391   1.0000   0.1645
  -3.500  -0.2573   0.02726   0.01684  -0.0384   1.0000   0.1794
  -3.250  -0.2325   0.02542   0.01480  -0.0377   1.0000   0.1969
  -3.000  -0.2074   0.02383   0.01306  -0.0370   1.0000   0.2230
  -2.750  -0.1822   0.02238   0.01164  -0.0362   1.0000   0.2578
  -2.500  -0.1562   0.02112   0.01042  -0.0355   1.0000   0.3064
  -2.250  -0.1304   0.01987   0.00952  -0.0350   1.0000   0.3733
  -2.000  -0.1071   0.01782   0.00867  -0.0336   1.0000   0.5404
  -1.750  -0.0707   0.01635   0.00779  -0.0326   1.0000   1.0000
  -1.500  -0.0493   0.01648   0.00748  -0.0318   1.0000   1.0000
  -1.250  -0.0281   0.01664   0.00731  -0.0311   1.0000   1.0000
  -1.000  -0.0069   0.01684   0.00720  -0.0305   1.0000   1.0000
  -0.750   0.0142   0.01707   0.00721  -0.0299   1.0000   1.0000
  -0.500   0.0351   0.01736   0.00729  -0.0294   1.0000   1.0000
  -0.250   0.0558   0.01769   0.00746  -0.0290   1.0000   1.0000
   0.000   0.0763   0.01808   0.00768  -0.0287   1.0000   1.0000
   0.250   0.0964   0.01853   0.00801  -0.0284   1.0000   1.0000
   0.500   0.1160   0.01904   0.00843  -0.0283   1.0000   1.0000
   0.750   0.1353   0.01963   0.00893  -0.0282   1.0000   1.0000
   1.000   0.1542   0.02029   0.00953  -0.0282   1.0000   1.0000
   1.250   0.1727   0.02102   0.01019  -0.0283   1.0000   1.0000
   1.500   0.1911   0.02181   0.01095  -0.0285   1.0000   1.0000
   1.750   0.2092   0.02266   0.01177  -0.0287   1.0000   1.0000
   2.000   0.2598   0.02378   0.01290  -0.0350   0.9857   1.0000
   2.250   0.3163   0.02480   0.01398  -0.0419   0.9660   1.0000
   2.500   0.3734   0.02565   0.01492  -0.0485   0.9451   1.0000
   2.750   0.4238   0.02632   0.01573  -0.0534   0.9227   1.0000
   3.000   0.4816   0.02677   0.01636  -0.0591   0.9012   1.0000
   3.250   0.5237   0.02719   0.01695  -0.0617   0.8776   1.0000
   3.500   0.5767   0.02729   0.01733  -0.0655   0.8557   1.0000
   3.750   0.6196   0.02730   0.01757  -0.0671   0.8318   1.0000
   4.000   0.6600   0.02700   0.01751  -0.0674   0.8048   1.0000
   4.250   0.7019   0.02598   0.01677  -0.0662   0.7741   1.0000
   4.500   0.7333   0.02512   0.01608  -0.0634   0.7398   1.0000
   4.750   0.7629   0.02442   0.01555  -0.0606   0.7059   1.0000
   5.000   0.7904   0.02396   0.01525  -0.0576   0.6697   1.0000
   5.250   0.8163   0.02373   0.01506  -0.0546   0.6283   1.0000
   5.500   0.8392   0.02383   0.01514  -0.0514   0.5787   1.0000
   5.750   0.8606   0.02424   0.01543  -0.0482   0.5237   1.0000
   6.000   0.8810   0.02489   0.01592  -0.0453   0.4671   1.0000
   6.250   0.8997   0.02549   0.01635  -0.0425   0.4122   1.0000
   6.500   0.9194   0.02641   0.01705  -0.0402   0.3645   1.0000
   6.750   0.9391   0.02767   0.01809  -0.0382   0.3205   1.0000
   7.000   0.9522   0.02884   0.01906  -0.0357   0.2676   1.0000
   7.250   0.9597   0.03015   0.02000  -0.0328   0.2079   1.0000
   7.500   0.9718   0.03305   0.02251  -0.0300   0.1490   1.0000
   7.750   0.9910   0.03603   0.02544  -0.0284   0.1194   1.0000
   8.000   1.0141   0.03931   0.02889  -0.0271   0.1061   1.0000
   8.250   1.0373   0.04271   0.03237  -0.0263   0.0983   1.0000
   8.500   1.0519   0.04639   0.03671  -0.0244   0.0939   1.0000
   8.750   1.0644   0.04980   0.04058  -0.0227   0.0891   1.0000
   9.000   1.0814   0.05346   0.04424  -0.0218   0.0850   1.0000
   9.250   1.0884   0.05814   0.04935  -0.0202   0.0845   1.0000
   9.500   1.0913   0.06292   0.05457  -0.0185   0.0846   1.0000
   9.750   1.0764   0.06745   0.05987  -0.0160   0.0862   1.0000
  10.000   1.0481   0.07313   0.06610  -0.0143   0.0885   1.0000
  10.250   1.0200   0.07844   0.07168  -0.0135   0.0903   1.0000
  10.500   0.9911   0.08415   0.07754  -0.0144   0.0921   1.0000
  10.750   0.9654   0.09081   0.08428  -0.0174   0.0938   1.0000
  11.000   0.9474   0.09790   0.09139  -0.0210   0.0956   1.0000
  11.250   0.9427   0.10410   0.09759  -0.0228   0.0970   1.0000
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