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NASA SC(2)-0402 AIRFOIL (sc20402-il) Xfoil prediction polar at RE=500,000 Ncrit=0


Details Polar file
Airfoil: NASA SC(2)-0402 AIRFOIL (sc20402-il)
Reynolds number: 500,000
Max Cl/Cd: 64.59 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20402-il-500000.txt
Download as CSV file: xf-sc20402-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0402 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
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.7221   0.13815   0.13593   0.0467   1.0000   0.0026
 -10.500  -0.7184   0.13430   0.13209   0.0455   1.0000   0.0027
 -10.250  -0.7147   0.13048   0.12828   0.0441   1.0000   0.0028
 -10.000  -0.7110   0.12666   0.12447   0.0427   1.0000   0.0030
  -9.750  -0.7072   0.12285   0.12067   0.0412   1.0000   0.0032
  -9.500  -0.7033   0.11903   0.11685   0.0396   1.0000   0.0035
  -9.250  -0.6994   0.11521   0.11305   0.0380   1.0000   0.0039
  -9.000  -0.6953   0.11138   0.10923   0.0362   1.0000   0.0043
  -8.750  -0.6906   0.10759   0.10546   0.0344   1.0000   0.0049
  -8.500  -0.6833   0.10414   0.10203   0.0321   1.0000   0.0056
  -5.750  -0.5137   0.05476   0.05188  -0.0161   1.0000   0.0065
  -5.500  -0.4922   0.04769   0.04463  -0.0201   1.0000   0.0067
  -5.250  -0.4689   0.04169   0.03840  -0.0233   1.0000   0.0071
  -5.000  -0.4448   0.03727   0.03378  -0.0256   1.0000   0.0077
  -4.750  -0.4186   0.03367   0.02998  -0.0273   1.0000   0.0086
  -4.500  -0.3903   0.03032   0.02639  -0.0287   1.0000   0.0097
  -3.500  -0.2682   0.01910   0.01388  -0.0313   1.0000   0.0193
  -3.250  -0.2393   0.01765   0.01222  -0.0316   1.0000   0.0261
  -3.000  -0.2094   0.01538   0.00966  -0.0321   1.0000   0.0329
  -2.750  -0.1802   0.01478   0.00887  -0.0319   1.0000   0.0408
  -2.000  -0.0920   0.00974   0.00342  -0.0306   1.0000   0.0180
  -1.750  -0.0600   0.00800   0.00157  -0.0309   1.0000   0.0089
  -1.500  -0.0310   0.00750   0.00094  -0.0309   1.0000   0.0084
  -1.250  -0.0005   0.00650   0.00047  -0.0318   1.0000   0.2319
  -1.000   0.0061   0.00411   0.00051  -0.0269   1.0000   0.9894
  -0.750   0.0335   0.00411   0.00046  -0.0269   1.0000   1.0000
  -0.500   0.0614   0.00411   0.00046  -0.0270   1.0000   1.0000
  -0.250   0.0891   0.00413   0.00048  -0.0271   1.0000   1.0000
   0.000   0.1340   0.00493   0.00048  -0.0304   0.7002   1.0000
   0.250   0.1554   0.00762   0.00092  -0.0301   0.0087   1.0000
   0.500   0.1839   0.00805   0.00150  -0.0300   0.0085   1.0000
   1.500   0.2993   0.01355   0.00745  -0.0269   0.0460   1.0000
   1.750   0.3274   0.01507   0.00917  -0.0267   0.0354   1.0000
   2.000   0.3383   0.00601   0.00061  -0.0242   0.0288   1.0000
   2.250   0.3688   0.00571   0.00059  -0.0234   0.0222   1.0000
   2.500   0.3913   0.00755   0.00259  -0.0236   0.0172   1.0000
   2.750   0.4145   0.01066   0.00620  -0.0229   0.0161   1.0000
   3.000   0.4452   0.01065   0.00644  -0.0217   0.0130   1.0000
   3.250   0.4706   0.01277   0.00885  -0.0211   0.0109   1.0000
   3.500   0.4947   0.01541   0.01174  -0.0208   0.0095   1.0000
   3.750   0.5176   0.01848   0.01505  -0.0207   0.0085   1.0000
   4.000   0.5384   0.02204   0.01880  -0.0210   0.0077   1.0000
   4.250   0.5569   0.02638   0.02333  -0.0215   0.0072   1.0000
   4.500   0.5766   0.03108   0.02850  -0.0220   0.0071   1.0000
   4.750   0.5944   0.03631   0.03395  -0.0228   0.0069   1.0000
   5.000   0.6085   0.04262   0.04042  -0.0242   0.0067   1.0000
   5.250   0.6188   0.05006   0.04800  -0.0261   0.0065   1.0000
   5.500   0.6901   0.06949   0.06733  -0.0342   0.0064   1.0000
   5.750   0.7084   0.07439   0.07234  -0.0370   0.0064   1.0000
   6.000   0.7254   0.07937   0.07740  -0.0402   0.0064   1.0000
   6.250   0.7410   0.08439   0.08248  -0.0439   0.0064   1.0000
   6.500   0.7554   0.08940   0.08754  -0.0481   0.0063   1.0000
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