Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NASA SC(2)-0402 AIRFOIL (sc20402-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0402 AIRFOIL (sc20402-il)
Reynolds number: 50,000
Max Cl/Cd: 14.49 at α=2.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20402-il-50000-n5.txt
Download as CSV file: xf-sc20402-il-50000-n5.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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.6954   0.11455   0.10811   0.0293   1.0000   0.0777
  -8.500  -0.6977   0.11211   0.10578   0.0243   1.0000   0.0801
  -8.250  -0.6962   0.10921   0.10296   0.0159   1.0000   0.0809
  -8.000  -0.6854   0.10297   0.09671   0.0233   1.0000   0.0855
  -7.750  -0.6792   0.09905   0.09282   0.0198   1.0000   0.0902
  -7.500  -0.6711   0.09572   0.08947   0.0049   1.0000   0.0940
  -7.000  -0.6494   0.08642   0.08014  -0.0008   1.0000   0.1072
  -6.000  -0.5445   0.05358   0.04700  -0.0064   1.0000   0.1618
  -5.500  -0.5252   0.05572   0.04769  -0.0182   1.0000   0.0637
  -5.250  -0.4929   0.05066   0.04217  -0.0206   1.0000   0.0385
  -5.000  -0.4652   0.04607   0.03720  -0.0229   1.0000   0.0327
  -4.750  -0.4329   0.04257   0.03308  -0.0244   1.0000   0.0274
  -4.500  -0.4058   0.03857   0.02872  -0.0259   1.0000   0.0253
  -4.250  -0.3760   0.03508   0.02473  -0.0270   1.0000   0.0233
  -4.000  -0.3455   0.03194   0.02105  -0.0277   1.0000   0.0216
  -3.750  -0.3149   0.02916   0.01772  -0.0279   1.0000   0.0201
  -3.500  -0.2849   0.02677   0.01480  -0.0277   1.0000   0.0189
  -3.250  -0.2564   0.02469   0.01232  -0.0272   1.0000   0.0187
  -3.000  -0.2296   0.02292   0.01030  -0.0267   1.0000   0.0217
  -2.750  -0.2036   0.02161   0.00870  -0.0257   1.0000   0.0262
  -2.500  -0.1775   0.02032   0.00717  -0.0249   1.0000   0.0301
  -2.250  -0.1506   0.01912   0.00581  -0.0246   1.0000   0.0340
  -2.000  -0.1230   0.01825   0.00474  -0.0245   1.0000   0.0367
  -1.250  -0.0561   0.01283   0.00168  -0.0186   1.0000   1.0000
  -1.000  -0.0289   0.01280   0.00138  -0.0185   1.0000   1.0000
  -0.750  -0.0018   0.01279   0.00118  -0.0185   1.0000   1.0000
  -0.500   0.0253   0.01279   0.00108  -0.0184   1.0000   1.0000
  -0.250   0.0522   0.01281   0.00108  -0.0184   1.0000   1.0000
   0.000   0.0790   0.01283   0.00118  -0.0183   1.0000   1.0000
   0.250   0.1058   0.01287   0.00137  -0.0181   1.0000   1.0000
   0.500   0.1326   0.01292   0.00167  -0.0180   1.0000   1.0000
   1.000   0.2083   0.01788   0.00429  -0.0205   0.0385   1.0000
   1.250   0.2362   0.01871   0.00533  -0.0202   0.0350   1.0000
   1.500   0.2638   0.01972   0.00652  -0.0199   0.0318   1.0000
   1.750   0.2915   0.02119   0.00818  -0.0196   0.0277   1.0000
   2.000   0.3203   0.02240   0.00972  -0.0191   0.0230   1.0000
   2.250   0.3486   0.02406   0.01163  -0.0187   0.0191   1.0000
   2.500   0.3771   0.02611   0.01403  -0.0182   0.0184   1.0000
   2.750   0.4066   0.02834   0.01675  -0.0176   0.0195   1.0000
   3.000   0.4360   0.03091   0.01985  -0.0171   0.0209   1.0000
   3.250   0.4648   0.03379   0.02326  -0.0167   0.0225   1.0000
   3.500   0.4928   0.03695   0.02693  -0.0166   0.0243   1.0000
   3.750   0.5191   0.04048   0.03089  -0.0168   0.0263   1.0000
   4.250   0.5747   0.04780   0.03925  -0.0175   0.0355   1.0000
   4.500   0.6019   0.05210   0.04397  -0.0182   0.0474   1.0000
   5.000   0.6799   0.06388   0.05699  -0.0262   0.1651   1.0000
   5.500   0.7106   0.07282   0.06682  -0.0278   0.1346   1.0000
   5.750   0.7244   0.07733   0.07143  -0.0296   0.1214   1.0000
   6.250   0.7494   0.08648   0.08080  -0.0380   0.1022   1.0000
   6.750   0.7686   0.09621   0.09056  -0.0410   0.0928   1.0000
   7.250   0.7843   0.10503   0.09943  -0.0471   0.0821   1.0000
   7.500   0.7903   0.11036   0.10471  -0.0487   0.0803   1.0000
   7.750   0.7896   0.11349   0.10779  -0.0568   0.0784   1.0000
   8.000   0.7900   0.11709   0.11133  -0.0605   0.0752   1.0000
   8.250   0.7931   0.12108   0.11529  -0.0623   0.0721   1.0000
   8.500   0.8087   0.12904   0.12328  -0.0564   0.0677   1.0000
   8.750   0.8071   0.13202   0.12619  -0.0611   0.0674   1.0000
   9.000   0.8052   0.13477   0.12886  -0.0654   0.0668   1.0000
<< Back to NASA SC(2)-0402 AIRFOIL (sc20402-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-0402 AIRFOIL (sc20402-il)