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NASA SC(2)-0403 AIRFOIL (sc20403-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0403 AIRFOIL (sc20403-il)
Reynolds number: 1,000,000
Max Cl/Cd: 46.54 at α=0.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20403-il-1000000.txt
Download as CSV file: xf-sc20403-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0403 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.6578   0.09432   0.09274   0.0200   1.0000   0.0047
  -7.750  -0.6536   0.09021   0.08864   0.0170   1.0000   0.0047
  -7.500  -0.6439   0.08532   0.08376   0.0115   1.0000   0.0047
  -7.250  -0.6306   0.07983   0.07826   0.0042   1.0000   0.0047
  -7.000  -0.6152   0.07422   0.07262  -0.0024   1.0000   0.0047
  -6.750  -0.5984   0.06877   0.06711  -0.0080   1.0000   0.0048
  -6.500  -0.5800   0.06346   0.06172  -0.0127   1.0000   0.0048
  -6.250  -0.5602   0.05830   0.05646  -0.0167   1.0000   0.0048
  -6.000  -0.5389   0.05330   0.05134  -0.0199   1.0000   0.0048
  -5.750  -0.5164   0.04848   0.04637  -0.0226   1.0000   0.0048
  -5.500  -0.4924   0.04385   0.04158  -0.0247   1.0000   0.0048
  -5.250  -0.4679   0.03563   0.03304  -0.0282   1.0000   0.0051
  -5.000  -0.4414   0.03008   0.02717  -0.0304   1.0000   0.0054
  -4.750  -0.4136   0.02639   0.02321  -0.0317   1.0000   0.0058
  -4.500  -0.3853   0.02351   0.02009  -0.0327   1.0000   0.0062
  -4.250  -0.3559   0.02086   0.01718  -0.0334   1.0000   0.0068
  -4.000  -0.3258   0.01843   0.01448  -0.0339   1.0000   0.0076
  -3.750  -0.2955   0.01634   0.01203  -0.0342   1.0000   0.0087
  -3.500  -0.2656   0.01475   0.01023  -0.0342   1.0000   0.0103
  -3.000  -0.2029   0.00950   0.00438  -0.0341   1.0000   0.0055
  -2.750  -0.1740   0.00873   0.00355  -0.0341   1.0000   0.0046
  -2.500  -0.1430   0.00744   0.00210  -0.0347   1.0000   0.0046
  -2.250  -0.1127   0.00668   0.00116  -0.0352   1.0000   0.0060
  -2.000  -0.0840   0.00633   0.00082  -0.0353   1.0000   0.0279
  -1.750  -0.0535   0.00531   0.00064  -0.0367   1.0000   0.2855
  -1.500  -0.0230   0.00396   0.00062  -0.0383   1.0000   0.6914
  -1.250   0.0040   0.00381   0.00061  -0.0381   1.0000   0.7443
  -1.000   0.0303   0.00368   0.00065  -0.0377   1.0000   0.8000
  -0.750   0.0563   0.00359   0.00069  -0.0372   1.0000   0.8354
  -0.500   0.0821   0.00354   0.00072  -0.0367   1.0000   0.8589
  -0.250   0.1081   0.00350   0.00075  -0.0363   1.0000   0.8777
   0.000   0.1336   0.00346   0.00080  -0.0357   1.0000   0.8953
   0.250   0.1587   0.00341   0.00085  -0.0350   1.0000   0.9133
   0.750   0.2175   0.00652   0.00101  -0.0365   0.0661   0.9456
   1.000   0.2417   0.00676   0.00110  -0.0355   0.0096   1.0000
   1.250   0.2706   0.00731   0.00178  -0.0356   0.0047   1.0000
   1.500   0.2990   0.00806   0.00265  -0.0355   0.0042   1.0000
   1.750   0.3263   0.00969   0.00443  -0.0350   0.0044   1.0000
   2.000   0.3551   0.01018   0.00497  -0.0348   0.0056   1.0000
   9.250   0.8113   0.13275   0.13135  -0.0803   0.0044   1.0000
   9.500   0.8145   0.13692   0.13552  -0.0820   0.0044   1.0000
   9.750   0.8179   0.14101   0.13960  -0.0837   0.0044   1.0000
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