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SOKOLOV AIRFOIL (sokolov-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: SOKOLOV AIRFOIL (sokolov-il)
Reynolds number: 500,000
Max Cl/Cd: 124.67 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sokolov-il-500000.txt
Download as CSV file: xf-sokolov-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: SOKOLOV 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -2.750   0.3467   0.01326   0.00769  -0.1502   0.8010   0.0410
  -2.500   0.3772   0.01235   0.00653  -0.1508   0.7891   0.0415
  -2.250   0.4066   0.01180   0.00582  -0.1511   0.7778   0.0433
  -2.000   0.4364   0.01119   0.00502  -0.1514   0.7669   0.0439
  -1.750   0.4658   0.01072   0.00442  -0.1517   0.7562   0.0446
  -1.500   0.4952   0.01034   0.00393  -0.1519   0.7454   0.0454
  -1.250   0.5244   0.01004   0.00353  -0.1521   0.7343   0.0461
  -1.000   0.5533   0.00982   0.00323  -0.1523   0.7231   0.0476
  -0.750   0.5822   0.00965   0.00296  -0.1524   0.7112   0.0489
  -0.500   0.6110   0.00950   0.00272  -0.1525   0.6978   0.0500
  -0.250   0.6400   0.00934   0.00247  -0.1526   0.6846   0.0523
   0.000   0.6689   0.00923   0.00231  -0.1528   0.6731   0.0568
   0.500   0.7260   0.00916   0.00229  -0.1530   0.6532   0.1097
   0.750   0.7544   0.00916   0.00229  -0.1531   0.6430   0.1311
   1.000   0.7827   0.00917   0.00231  -0.1532   0.6336   0.1450
   1.250   0.8109   0.00919   0.00234  -0.1533   0.6246   0.1675
   1.500   0.8397   0.00901   0.00247  -0.1537   0.6151   0.2951
   1.750   0.8680   0.00894   0.00263  -0.1539   0.6060   0.4130
   2.000   0.8961   0.00886   0.00276  -0.1541   0.5975   0.5175
   2.750   0.9746   0.00830   0.00296  -0.1529   0.5746   1.0000
   4.500   1.1664   0.00937   0.00387  -0.1523   0.4976   1.0000
   4.750   1.1931   0.00957   0.00403  -0.1521   0.4790   1.0000
   5.000   1.2191   0.00984   0.00422  -0.1519   0.4454   1.0000
   5.250   1.2413   0.01060   0.00458  -0.1511   0.3681   1.0000
   5.500   1.2637   0.01140   0.00509  -0.1504   0.3182   1.0000
   5.750   1.2859   0.01221   0.00563  -0.1497   0.2678   1.0000
   6.000   1.3002   0.01407   0.00672  -0.1479   0.1387   1.0000
   6.250   1.3121   0.01614   0.00806  -0.1457   0.0244   1.0000
   6.500   1.3348   0.01675   0.00874  -0.1448   0.0209   1.0000
   6.750   1.3560   0.01751   0.00959  -0.1436   0.0184   1.0000
   7.000   1.3760   0.01837   0.01056  -0.1422   0.0170   1.0000
   7.250   1.3964   0.01910   0.01136  -0.1410   0.0159   1.0000
   7.500   1.4150   0.01998   0.01233  -0.1394   0.0150   1.0000
   7.750   1.4317   0.02096   0.01340  -0.1376   0.0143   1.0000
   8.000   1.4464   0.02207   0.01460  -0.1355   0.0137   1.0000
   8.250   1.4579   0.02336   0.01597  -0.1328   0.0132   1.0000
   8.500   1.4648   0.02490   0.01761  -0.1295   0.0127   1.0000
   8.750   1.4674   0.02674   0.01953  -0.1255   0.0124   1.0000
   9.000   1.4659   0.02906   0.02195  -0.1209   0.0121   1.0000
   9.250   1.4759   0.03030   0.02329  -0.1182   0.0117   1.0000
   9.500   1.4872   0.03136   0.02445  -0.1158   0.0113   1.0000
   9.750   1.4974   0.03290   0.02611  -0.1134   0.0111   1.0000
  10.000   1.5084   0.03460   0.02792  -0.1111   0.0109   1.0000
  10.250   1.5205   0.03646   0.02990  -0.1089   0.0107   1.0000
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