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)-0406 AIRFOIL (sc20406-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0406 AIRFOIL (sc20406-il)
Reynolds number: 100,000
Max Cl/Cd: 37.15 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20406-il-100000.txt
Download as CSV file: xf-sc20406-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0406 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5493   0.08747   0.08284  -0.0001   1.0000   0.1071
  -8.500  -0.5672   0.08273   0.07818  -0.0050   1.0000   0.1105
  -8.250  -0.5935   0.07704   0.07255  -0.0133   1.0000   0.1113
  -8.000  -0.6639   0.08414   0.07943  -0.0016   1.0000   0.1042
  -7.500  -0.6602   0.07231   0.06746  -0.0155   1.0000   0.1143
  -7.250  -0.6480   0.06864   0.06378  -0.0158   1.0000   0.1206
  -7.000  -0.6396   0.06339   0.05838  -0.0201   1.0000   0.1291
  -6.750  -0.6275   0.05890   0.05369  -0.0234   1.0000   0.1416
  -6.500  -0.6123   0.05496   0.04962  -0.0250   1.0000   0.1553
  -6.250  -0.5954   0.05130   0.04585  -0.0256   1.0000   0.1702
  -6.000  -0.5469   0.03809   0.03070  -0.0322   1.0000   0.0694
  -5.750  -0.5217   0.03330   0.02552  -0.0328   1.0000   0.0644
  -5.500  -0.4938   0.02995   0.02165  -0.0332   1.0000   0.0650
  -5.250  -0.4647   0.02662   0.01775  -0.0332   1.0000   0.0636
  -5.000  -0.4358   0.02391   0.01458  -0.0329   1.0000   0.0640
  -4.750  -0.4074   0.02175   0.01208  -0.0324   1.0000   0.0675
  -4.500  -0.3804   0.01984   0.01008  -0.0321   1.0000   0.0778
  -4.250  -0.3535   0.01818   0.00839  -0.0316   1.0000   0.0949
  -4.000  -0.3274   0.01663   0.00696  -0.0314   1.0000   0.1186
  -3.750  -0.3008   0.01526   0.00576  -0.0312   1.0000   0.1498
  -3.500  -0.2758   0.01200   0.00462  -0.0320   1.0000   0.4988
  -3.250  -0.2641   0.01178   0.00511  -0.0262   1.0000   0.7328
  -3.000  -0.2496   0.01189   0.00523  -0.0217   1.0000   0.7939
  -2.750  -0.2377   0.01192   0.00520  -0.0168   1.0000   0.8358
  -2.500  -0.2280   0.01182   0.00508  -0.0114   1.0000   0.8721
  -2.250  -0.2178   0.01158   0.00480  -0.0064   1.0000   0.9080
  -2.000  -0.1977   0.01120   0.00436  -0.0030   1.0000   0.9583
  -1.750  -0.1019   0.01112   0.00393  -0.0158   1.0000   1.0000
  -1.500  -0.0879   0.01083   0.00360  -0.0144   1.0000   1.0000
  -1.250  -0.0759   0.01058   0.00332  -0.0125   1.0000   1.0000
  -1.000  -0.0591   0.01041   0.00312  -0.0115   1.0000   1.0000
  -0.750  -0.0355   0.01032   0.00298  -0.0117   1.0000   1.0000
  -0.500  -0.0087   0.01029   0.00290  -0.0124   1.0000   1.0000
  -0.250   0.0194   0.01030   0.00288  -0.0132   1.0000   1.0000
   0.000   0.0479   0.01034   0.00291  -0.0140   1.0000   1.0000
   0.250   0.0765   0.01040   0.00298  -0.0148   1.0000   1.0000
   0.500   0.1049   0.01049   0.00312  -0.0155   1.0000   1.0000
   0.750   0.1332   0.01061   0.00328  -0.0162   1.0000   1.0000
   1.000   0.1613   0.01075   0.00349  -0.0168   1.0000   1.0000
   1.250   0.1891   0.01091   0.00376  -0.0174   1.0000   1.0000
   1.500   0.2167   0.01110   0.00407  -0.0179   1.0000   1.0000
   1.750   0.2440   0.01133   0.00448  -0.0184   1.0000   1.0000
   2.000   0.2710   0.01158   0.00492  -0.0189   1.0000   1.0000
   2.250   0.3033   0.01185   0.00543  -0.0204   0.9963   1.0000
   2.500   0.4068   0.01095   0.00491  -0.0304   0.7760   1.0000
   2.750   0.4147   0.01480   0.00555  -0.0252   0.1711   1.0000
   3.000   0.4400   0.01630   0.00677  -0.0247   0.1287   1.0000
   3.250   0.4662   0.01781   0.00812  -0.0243   0.1024   1.0000
   3.500   0.4929   0.01951   0.00965  -0.0241   0.0813   1.0000
   3.750   0.5217   0.02130   0.01164  -0.0237   0.0707   1.0000
   4.000   0.5501   0.02338   0.01394  -0.0233   0.0628   1.0000
   4.250   0.5792   0.02585   0.01683  -0.0228   0.0610   1.0000
   4.500   0.6075   0.02892   0.02041  -0.0221   0.0620   1.0000
   4.750   0.6341   0.03269   0.02462  -0.0216   0.0650   1.0000
   5.000   0.6613   0.03809   0.03073  -0.0206   0.0789   1.0000
   5.500   0.7225   0.05282   0.04751  -0.0216   0.1647   1.0000
   5.750   0.7364   0.05635   0.05130  -0.0224   0.1422   1.0000
   7.000   0.7888   0.07903   0.07472  -0.0302   0.1026   1.0000
   7.250   0.7956   0.08341   0.07917  -0.0319   0.0977   1.0000
   7.500   0.8182   0.08839   0.08387  -0.0260   0.0942   1.0000
   7.750   0.8047   0.09367   0.08943  -0.0338   0.0929   1.0000
   8.000   0.7913   0.09954   0.09535  -0.0443   0.0905   1.0000
   8.250   0.7868   0.10453   0.10030  -0.0493   0.0876   1.0000
   8.500   0.7880   0.10886   0.10460  -0.0513   0.0847   1.0000
   8.750   0.6758   0.10449   0.10043  -0.0320   0.0934   1.0000
   9.000   0.6565   0.10909   0.10497  -0.0372   0.0927   1.0000
<< Back to NASA SC(2)-0406 AIRFOIL (sc20406-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-0406 AIRFOIL (sc20406-il)