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NASA SC(2)-0402 AIRFOIL (sc20402-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0402 AIRFOIL (sc20402-il)
Reynolds number: 1,000,000
Max Cl/Cd: 32.69 at α=-0.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20402-il-1000000.txt
Download as CSV file: xf-sc20402-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.5843   0.13426   0.13271   0.0303   1.0000   0.0019
 -11.000  -0.5830   0.13064   0.12910   0.0294   1.0000   0.0019
  -7.250  -0.6485   0.07947   0.07799   0.0098   1.0000   0.0023
  -7.000  -0.6381   0.07520   0.07370   0.0062   1.0000   0.0026
  -6.750  -0.6209   0.07026   0.06872   0.0005   1.0000   0.0028
  -6.500  -0.6007   0.06503   0.06343  -0.0052   1.0000   0.0029
  -6.250  -0.5785   0.05982   0.05813  -0.0104   1.0000   0.0029
  -6.000  -0.5544   0.05471   0.05292  -0.0150   1.0000   0.0030
  -5.750  -0.5288   0.04971   0.04778  -0.0188   1.0000   0.0030
  -5.500  -0.5012   0.04467   0.04255  -0.0221   1.0000   0.0028
  -5.250  -0.4680   0.03917   0.03664  -0.0245   1.0000   0.0023
  -5.000  -0.4404   0.03498   0.03224  -0.0263   1.0000   0.0024
  -4.750  -0.4123   0.03121   0.02826  -0.0278   1.0000   0.0025
  -4.500  -0.3835   0.02779   0.02461  -0.0291   1.0000   0.0026
  -4.250  -0.3540   0.02466   0.02124  -0.0300   1.0000   0.0028
  -4.000  -0.3240   0.02183   0.01816  -0.0307   1.0000   0.0030
  -3.750  -0.2936   0.01930   0.01536  -0.0312   1.0000   0.0034
  -3.500  -0.2632   0.01705   0.01286  -0.0314   1.0000   0.0038
  -3.250  -0.2330   0.01524   0.01083  -0.0315   1.0000   0.0046
  -3.000  -0.2046   0.01530   0.01079  -0.0308   1.0000   0.0060
  -2.750  -0.1738   0.01188   0.00704  -0.0320   1.0000   0.0099
  -1.250   0.0031   0.00602   0.00053  -0.0321   1.0000   0.0057
  -1.000   0.0360   0.00395   0.00036  -0.0346   1.0000   0.6490
  -0.500   0.0707   0.00296   0.00041  -0.0292   1.0000   1.0000
  -0.250   0.1131   0.00346   0.00037  -0.0323   0.7973   1.0000
   0.000   0.1340   0.00624   0.00064  -0.0320   0.0057   1.0000
   0.250   0.1626   0.00662   0.00117  -0.0319   0.0056   1.0000
   0.500   0.1912   0.00709   0.00174  -0.0318   0.0079   1.0000
   1.750   0.3304   0.01447   0.00965  -0.0300   0.0062   1.0000
   2.000   0.3595   0.01520   0.01057  -0.0293   0.0048   1.0000
   2.250   0.3880   0.01682   0.01239  -0.0288   0.0039   1.0000
   2.500   0.4162   0.01884   0.01466  -0.0282   0.0034   1.0000
   2.750   0.4439   0.02114   0.01720  -0.0277   0.0031   1.0000
   3.000   0.4711   0.02367   0.01998  -0.0272   0.0028   1.0000
   3.250   0.4976   0.02644   0.02299  -0.0269   0.0026   1.0000
   3.500   0.5235   0.02950   0.02628  -0.0267   0.0025   1.0000
   3.750   0.5488   0.03287   0.02987  -0.0266   0.0024   1.0000
   4.000   0.5737   0.03670   0.03393  -0.0267   0.0023   1.0000
   4.250   0.6026   0.04116   0.03879  -0.0266   0.0026   1.0000
   4.500   0.6291   0.04591   0.04379  -0.0273   0.0029   1.0000
   4.750   0.6534   0.05045   0.04849  -0.0286   0.0029   1.0000
   5.000   0.6765   0.05509   0.05328  -0.0304   0.0028   1.0000
   5.250   0.6985   0.05987   0.05818  -0.0328   0.0028   1.0000
   5.500   0.7190   0.06477   0.06319  -0.0357   0.0027   1.0000
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