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GOE 590 AIRFOIL (goe590-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 590 AIRFOIL (goe590-il)
Reynolds number: 50,000
Max Cl/Cd: 38.37 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe590-il-50000-n5.txt
Download as CSV file: xf-goe590-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 590 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4015   0.10396   0.09745  -0.0132   1.0000   0.0609
  -7.750  -0.4022   0.10225   0.09586  -0.0157   1.0000   0.0620
  -7.500  -0.4004   0.10076   0.09444  -0.0198   1.0000   0.0627
  -7.250  -0.3946   0.09904   0.09276  -0.0242   1.0000   0.0631
  -7.000  -0.3842   0.09151   0.08531  -0.0189   1.0000   0.0667
  -6.750  -0.3767   0.08828   0.08212  -0.0199   1.0000   0.0699
  -6.500  -0.3689   0.08575   0.07963  -0.0229   1.0000   0.0745
  -6.250  -0.3570   0.08476   0.07859  -0.0290   1.0000   0.0766
  -6.000  -0.3487   0.07996   0.07388  -0.0285   1.0000   0.0780
  -5.750  -0.3405   0.07572   0.06966  -0.0271   1.0000   0.0812
  -5.500  -0.3293   0.07262   0.06656  -0.0281   1.0000   0.0853
  -5.250  -0.3085   0.07199   0.06570  -0.0337   1.0000   0.0906
  -5.000  -0.3017   0.06699   0.06086  -0.0321   1.0000   0.0922
  -4.750  -0.2925   0.06345   0.05737  -0.0311   1.0000   0.0953
  -4.500  -0.2799   0.06060   0.05449  -0.0312   1.0000   0.0991
  -4.000  -0.2562   0.05536   0.04915  -0.0312   1.0000   0.1085
  -3.500  -0.1964   0.04775   0.04102  -0.0361   0.9849   0.0693
  -3.250  -0.1529   0.04321   0.03604  -0.0404   0.9693   0.0552
  -3.000  -0.1049   0.03923   0.03127  -0.0441   0.9549   0.0492
  -2.750  -0.0673   0.03554   0.02722  -0.0470   0.9404   0.0512
  -2.500  -0.0305   0.03295   0.02427  -0.0493   0.9247   0.0579
  -2.250   0.0085   0.03001   0.02070  -0.0510   0.9095   0.0591
  -2.000   0.0452   0.02784   0.01804  -0.0524   0.8935   0.0685
  -1.750   0.0831   0.02573   0.01525  -0.0535   0.8779   0.0740
  -1.500   0.1209   0.02385   0.01280  -0.0546   0.8625   0.0788
  -1.250   0.1586   0.02278   0.01139  -0.0560   0.8464   0.0946
  -1.000   0.1966   0.02169   0.00987  -0.0571   0.8308   0.1016
  -0.750   0.2273   0.02089   0.00877  -0.0570   0.8142   0.1084
  -0.500   0.2546   0.02033   0.00798  -0.0564   0.7983   0.1192
  -0.250   0.2799   0.01980   0.00738  -0.0557   0.7833   0.1396
   0.000   0.3493   0.01686   0.00642  -0.0640   0.7711   1.0000
   0.250   0.3740   0.01707   0.00621  -0.0631   0.7570   1.0000
   0.500   0.3985   0.01731   0.00613  -0.0623   0.7440   1.0000
   0.750   0.4229   0.01756   0.00611  -0.0614   0.7320   1.0000
   1.000   0.4469   0.01784   0.00621  -0.0607   0.7200   1.0000
   1.250   0.4709   0.01815   0.00637  -0.0600   0.7087   1.0000
   1.500   0.4949   0.01846   0.00656  -0.0593   0.6984   1.0000
   1.750   0.5195   0.01876   0.00676  -0.0586   0.6895   1.0000
   2.000   0.5431   0.01912   0.00713  -0.0581   0.6795   1.0000
   2.250   0.5674   0.01949   0.00748  -0.0575   0.6708   1.0000
   2.500   0.5923   0.01984   0.00784  -0.0570   0.6628   1.0000
   2.750   0.6163   0.02028   0.00840  -0.0566   0.6543   1.0000
   3.000   0.6413   0.02064   0.00881  -0.0561   0.6476   1.0000
   3.250   0.6647   0.02115   0.00946  -0.0557   0.6393   1.0000
   3.500   0.6897   0.02155   0.00996  -0.0552   0.6334   1.0000
   3.750   0.7126   0.02214   0.01083  -0.0547   0.6255   1.0000
   4.000   0.7374   0.02259   0.01146  -0.0542   0.6198   1.0000
   4.250   0.7599   0.02326   0.01241  -0.0537   0.6124   1.0000
   4.500   0.7842   0.02379   0.01321  -0.0532   0.6068   1.0000
   4.750   0.8067   0.02452   0.01430  -0.0527   0.6002   1.0000
   5.000   0.8308   0.02494   0.01513  -0.0517   0.5921   1.0000
   5.250   0.8497   0.02354   0.01378  -0.0468   0.5475   1.0000
   5.500   0.8575   0.02235   0.01242  -0.0404   0.4600   1.0000
   5.750   0.8461   0.02479   0.01272  -0.0346   0.1096   1.0000
   6.000   0.8505   0.02781   0.01535  -0.0323   0.0690   1.0000
   6.250   0.8586   0.02992   0.01760  -0.0300   0.0556   1.0000
   6.500   0.8654   0.03193   0.01975  -0.0277   0.0479   1.0000
   6.750   0.8713   0.03393   0.02201  -0.0250   0.0444   1.0000
   7.000   0.8805   0.03564   0.02391  -0.0226   0.0403   1.0000
   7.250   0.8909   0.03769   0.02594  -0.0202   0.0359   1.0000
   7.500   0.9209   0.03944   0.02797  -0.0185   0.0336   1.0000
   7.750   0.9587   0.04195   0.03072  -0.0179   0.0312   1.0000
   8.000   0.9866   0.04509   0.03394  -0.0176   0.0280   1.0000
   8.250   1.0091   0.04876   0.03793  -0.0167   0.0268   1.0000
   8.500   1.0261   0.05223   0.04185  -0.0152   0.0266   1.0000
   8.750   1.0388   0.05586   0.04590  -0.0136   0.0266   1.0000
   9.000   1.0472   0.05967   0.05023  -0.0120   0.0267   1.0000
   9.250   1.0521   0.06352   0.05442  -0.0104   0.0269   1.0000
   9.500   1.0541   0.06755   0.05876  -0.0088   0.0270   1.0000
   9.750   1.0528   0.07168   0.06316  -0.0074   0.0272   1.0000
  10.750   0.9938   0.08454   0.07721  -0.0025   0.0287   1.0000
  11.000   0.9752   0.08917   0.08203  -0.0043   0.0290   1.0000
  11.250   0.9560   0.09456   0.08757  -0.0072   0.0295   1.0000
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