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NASA SC(2)-0406 AIRFOIL (sc20406-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0406 AIRFOIL (sc20406-il)
Reynolds number: 500,000
Max Cl/Cd: 50.28 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20406-il-500000-n5.txt
Download as CSV file: xf-sc20406-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0406 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.5737   0.10795   0.10575   0.0091   1.0000   0.0074
 -10.750  -0.5785   0.10245   0.10026   0.0076   1.0000   0.0072
  -9.750  -0.7048   0.09535   0.09311   0.0140   1.0000   0.0069
  -9.500  -0.7100   0.08944   0.08723   0.0105   1.0000   0.0068
  -9.250  -0.7155   0.08313   0.08095   0.0062   1.0000   0.0068
  -9.000  -0.7232   0.07589   0.07375  -0.0002   1.0000   0.0067
  -8.750  -0.7293   0.06549   0.06329  -0.0136   1.0000   0.0066
  -8.500  -0.7319   0.05622   0.05382  -0.0206   1.0000   0.0066
  -8.250  -0.7312   0.04623   0.04348  -0.0252   1.0000   0.0066
  -8.000  -0.7281   0.03371   0.03026  -0.0282   1.0000   0.0068
  -7.750  -0.7097   0.02860   0.02459  -0.0292   1.0000   0.0069
  -7.500  -0.6899   0.02284   0.01807  -0.0303   1.0000   0.0072
  -7.250  -0.6654   0.01990   0.01468  -0.0309   1.0000   0.0074
  -7.000  -0.6395   0.01830   0.01286  -0.0312   1.0000   0.0079
  -6.750  -0.6129   0.01720   0.01156  -0.0315   1.0000   0.0083
  -6.500  -0.5860   0.01579   0.00994  -0.0317   1.0000   0.0084
  -6.250  -0.5588   0.01452   0.00847  -0.0318   1.0000   0.0084
  -6.000  -0.5313   0.01349   0.00729  -0.0320   1.0000   0.0085
  -5.750  -0.5035   0.01262   0.00630  -0.0323   1.0000   0.0087
  -5.500  -0.4755   0.01189   0.00547  -0.0325   1.0000   0.0090
  -5.250  -0.4472   0.01127   0.00474  -0.0328   1.0000   0.0093
  -5.000  -0.4189   0.01075   0.00415  -0.0331   1.0000   0.0099
  -4.750  -0.3905   0.01031   0.00364  -0.0334   1.0000   0.0107
  -4.500  -0.3620   0.00992   0.00319  -0.0336   1.0000   0.0119
  -4.250  -0.3336   0.00955   0.00282  -0.0338   1.0000   0.0158
  -4.000  -0.3054   0.00924   0.00252  -0.0341   1.0000   0.0247
  -3.750  -0.2772   0.00897   0.00231  -0.0343   1.0000   0.0367
  -3.500  -0.2492   0.00874   0.00211  -0.0345   1.0000   0.0466
  -3.250  -0.2213   0.00853   0.00194  -0.0346   1.0000   0.0573
  -3.000  -0.1934   0.00827   0.00178  -0.0348   1.0000   0.0820
  -2.750  -0.1652   0.00785   0.00159  -0.0352   1.0000   0.1406
  -2.500  -0.1369   0.00740   0.00144  -0.0357   1.0000   0.2225
  -2.250  -0.1084   0.00689   0.00129  -0.0363   1.0000   0.3289
  -2.000  -0.0794   0.00627   0.00120  -0.0371   1.0000   0.4765
  -1.750  -0.0515   0.00595   0.00120  -0.0374   1.0000   0.5708
  -1.500  -0.0243   0.00578   0.00124  -0.0373   1.0000   0.6285
  -1.250   0.0026   0.00570   0.00129  -0.0371   1.0000   0.6675
  -1.000   0.0368   0.00561   0.00130  -0.0386   0.9957   0.6941
  -0.750   0.0781   0.00551   0.00128  -0.0415   0.9814   0.7175
  -0.500   0.1158   0.00545   0.00129  -0.0435   0.9576   0.7416
  -0.250   0.1464   0.00545   0.00133  -0.0437   0.9268   0.7605
   0.000   0.1705   0.00556   0.00134  -0.0423   0.8718   0.7727
   0.250   0.1947   0.00576   0.00135  -0.0411   0.8081   0.7832
   0.500   0.2191   0.00616   0.00136  -0.0401   0.7036   0.7931
   0.750   0.2434   0.00697   0.00145  -0.0395   0.5121   0.8013
   1.000   0.2693   0.00778   0.00162  -0.0395   0.3317   0.8098
   1.250   0.2958   0.00844   0.00181  -0.0396   0.1940   0.8187
   1.500   0.3226   0.00897   0.00198  -0.0397   0.0990   0.8274
   1.750   0.3502   0.00923   0.00214  -0.0397   0.0672   0.8364
   2.000   0.3775   0.00941   0.00231  -0.0396   0.0537   0.8457
   2.250   0.4046   0.00959   0.00252  -0.0394   0.0421   0.8559
   2.500   0.4316   0.00981   0.00272  -0.0392   0.0304   0.8665
   2.750   0.4585   0.01005   0.00296  -0.0389   0.0198   0.8768
   3.000   0.4850   0.01034   0.00326  -0.0386   0.0137   0.8870
   3.250   0.5110   0.01061   0.00360  -0.0381   0.0125   0.8981
   3.500   0.5366   0.01091   0.00402  -0.0375   0.0117   0.9109
   3.750   0.5612   0.01124   0.00443  -0.0367   0.0111   0.9261
   4.000   0.5843   0.01162   0.00490  -0.0355   0.0101   0.9463
   4.250   0.6093   0.01229   0.00570  -0.0349   0.0093   1.0000
   4.500   0.6368   0.01295   0.00643  -0.0349   0.0091   1.0000
   4.750   0.6641   0.01366   0.00727  -0.0349   0.0089   1.0000
   5.000   0.6911   0.01449   0.00821  -0.0348   0.0086   1.0000
   5.250   0.7177   0.01550   0.00938  -0.0346   0.0082   1.0000
   5.500   0.7439   0.01674   0.01081  -0.0343   0.0079   1.0000
   5.750   0.7696   0.01831   0.01263  -0.0339   0.0076   1.0000
   6.000   0.7944   0.02035   0.01500  -0.0334   0.0075   1.0000
   6.250   0.8174   0.02324   0.01835  -0.0327   0.0074   1.0000
   6.500   0.8367   0.02772   0.02344  -0.0316   0.0074   1.0000
   6.750   0.8539   0.03250   0.02877  -0.0306   0.0070   1.0000
   7.000   0.8686   0.03779   0.03449  -0.0298   0.0066   1.0000
   7.250   0.8723   0.04809   0.04536  -0.0290   0.0065   1.0000
   7.500   0.8769   0.05612   0.05373  -0.0296   0.0064   1.0000
   7.750   0.8790   0.06376   0.06160  -0.0313   0.0064   1.0000
   8.000   0.8778   0.07137   0.06938  -0.0344   0.0063   1.0000
   8.250   0.8711   0.07989   0.07799  -0.0397   0.0064   1.0000
   8.500   0.8585   0.08817   0.08631  -0.0475   0.0064   1.0000
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