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GOE 620 AIRFOIL (goe620-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 620 AIRFOIL (goe620-il)
Reynolds number: 50,000
Max Cl/Cd: 14.33 at α=7.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe620-il-50000-n5.txt
Download as CSV file: xf-goe620-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 620 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.1552   0.11882   0.11136  -0.0777   0.9404   0.0970
 -10.000  -0.1485   0.11484   0.10736  -0.0807   0.9360   0.0960
  -9.750  -0.1497   0.11173   0.10427  -0.0817   0.9286   0.0960
  -9.500  -0.1439   0.10793   0.10046  -0.0846   0.9238   0.0964
  -9.250  -0.1402   0.10414   0.09665  -0.0874   0.9190   0.0967
  -9.000  -0.1465   0.10104   0.09358  -0.0878   0.9109   0.0967
  -8.750  -0.1451   0.09673   0.08926  -0.0909   0.9059   0.0966
  -8.500  -0.1547   0.09313   0.08568  -0.0919   0.8985   0.0962
  -8.250  -0.1665   0.08909   0.08167  -0.0932   0.8909   0.0958
  -8.000  -0.1761   0.08360   0.07617  -0.0972   0.8857   0.0960
  -7.750  -0.2177   0.07990   0.07253  -0.0951   0.8735   0.0955
  -7.500  -0.2600   0.07072   0.06319  -0.1002   0.8642   0.0953
  -7.250  -0.2903   0.06495   0.05719  -0.1008   0.8544   0.0955
  -7.000  -0.2924   0.05914   0.05094  -0.1041   0.8496   0.0972
  -6.750  -0.3088   0.05645   0.04797  -0.1010   0.8394   0.0980
  -6.500  -0.2999   0.05249   0.04346  -0.1020   0.8342   0.0997
  -6.250  -0.2855   0.04928   0.03964  -0.1026   0.8296   0.1013
  -6.000  -0.2800   0.04853   0.03888  -0.0999   0.8213   0.1027
  -5.750  -0.2560   0.04744   0.03769  -0.1001   0.8165   0.1055
  -5.500  -0.2268   0.04578   0.03573  -0.1012   0.8132   0.1093
  -5.250  -0.2245   0.04498   0.03466  -0.0980   0.8041   0.1116
  -5.000  -0.2005   0.04378   0.03326  -0.0979   0.7989   0.1145
  -4.750  -0.1694   0.04288   0.03230  -0.0987   0.7954   0.1187
  -4.500  -0.1615   0.04256   0.03184  -0.0961   0.7870   0.1226
  -4.250  -0.1384   0.04182   0.03094  -0.0956   0.7813   0.1278
  -4.000  -0.1070   0.04107   0.03014  -0.0963   0.7775   0.1343
  -3.750  -0.0948   0.04085   0.02977  -0.0942   0.7698   0.1407
  -3.500  -0.0742   0.04058   0.02955  -0.0932   0.7632   0.1488
  -3.250  -0.0428   0.03996   0.02890  -0.0938   0.7593   0.1617
  -3.000  -0.0274   0.03996   0.02887  -0.0921   0.7520   0.1739
  -2.750  -0.0079   0.03989   0.02880  -0.0911   0.7449   0.1901
  -2.500   0.0241   0.03953   0.02843  -0.0917   0.7408   0.2136
  -2.250   0.0401   0.03976   0.02865  -0.0901   0.7335   0.2326
  -2.000   0.0572   0.03999   0.02888  -0.0887   0.7260   0.2532
  -1.750   0.0882   0.03988   0.02878  -0.0890   0.7219   0.2832
  -1.500   0.1006   0.04042   0.02935  -0.0870   0.7138   0.3048
  -1.250   0.1202   0.04070   0.02963  -0.0859   0.7067   0.3272
  -1.000   0.1534   0.04050   0.02937  -0.0864   0.7029   0.3509
  -0.750   0.1648   0.04112   0.02997  -0.0845   0.6941   0.3667
  -0.500   0.1877   0.04132   0.03014  -0.0839   0.6876   0.3869
  -0.250   0.2215   0.04110   0.02989  -0.0845   0.6840   0.4122
   0.000   0.2291   0.04196   0.03078  -0.0822   0.6744   0.4291
   0.250   0.2547   0.04207   0.03090  -0.0819   0.6685   0.4541
   0.500   0.2891   0.04178   0.03065  -0.0825   0.6652   0.4839
   0.750   0.2927   0.04292   0.03184  -0.0800   0.6545   0.5038
   1.000   0.3211   0.04286   0.03186  -0.0799   0.6495   0.5375
   1.500   0.3561   0.04380   0.03306  -0.0777   0.6350   0.6131
   1.750   0.3877   0.04333   0.03291  -0.0777   0.6308   0.6885
   2.000   0.4487   0.04243   0.03238  -0.0829   0.6281   1.0000
   2.250   0.4452   0.04434   0.03416  -0.0803   0.6157   1.0000
   2.500   0.4791   0.04426   0.03383  -0.0808   0.6119   1.0000
   2.750   0.4809   0.04595   0.03541  -0.0787   0.6008   1.0000
   3.000   0.5100   0.04607   0.03535  -0.0787   0.5960   1.0000
   3.250   0.5459   0.04573   0.03483  -0.0791   0.5930   1.0000
   3.500   0.5421   0.04781   0.03686  -0.0766   0.5803   1.0000
   3.750   0.5759   0.04752   0.03642  -0.0768   0.5767   1.0000
   4.000   0.5758   0.04945   0.03830  -0.0746   0.5648   1.0000
   4.250   0.6069   0.04926   0.03798  -0.0745   0.5604   1.0000
   4.750   0.6385   0.05094   0.03952  -0.0723   0.5440   1.0000
   5.000   0.6740   0.05028   0.03876  -0.0723   0.5410   1.0000
   5.250   0.6694   0.05274   0.04120  -0.0702   0.5278   1.0000
   5.500   0.7024   0.05221   0.04058  -0.0700   0.5242   1.0000
   6.000   0.7291   0.05437   0.04268  -0.0677   0.5074   1.0000
   6.250   0.7641   0.05359   0.04183  -0.0675   0.5049   1.0000
   6.500   0.7529   0.05697   0.04523  -0.0655   0.4908   1.0000
   6.750   0.7856   0.05632   0.04453  -0.0651   0.4878   1.0000
   7.250   0.8048   0.05944   0.04764  -0.0629   0.4707   1.0000
   7.500   0.8395   0.05857   0.04673  -0.0626   0.4684   1.0000
   8.000   0.8334   0.06438   0.05260  -0.0600   0.4464   1.0000
   8.250   0.8353   0.06700   0.05524  -0.0590   0.4369   1.0000
   8.500   0.8633   0.06676   0.05499  -0.0585   0.4340   1.0000
   9.000   0.8699   0.07183   0.06012  -0.0569   0.4171   1.0000
   9.250   0.9002   0.07123   0.05950  -0.0563   0.4149   1.0000
   9.750   0.9005   0.07726   0.06562  -0.0550   0.3980   1.0000
  11.750   0.9032   0.10235   0.09108  -0.0529   0.3363   1.0000
  12.000   0.9288   0.10171   0.09045  -0.0519   0.3317   1.0000
  12.250   0.9624   0.09981   0.08858  -0.0507   0.3290   1.0000
  12.750   0.9645   0.10570   0.09458  -0.0506   0.3131   1.0000
  13.250   0.9608   0.11290   0.10192  -0.0513   0.2978   1.0000
  13.750   0.9595   0.11984   0.10902  -0.0522   0.2825   1.0000
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