Airfoil Tools
Search 1638 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NASA SC(2)-0012 AIRFOIL (sc20012-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0012 AIRFOIL (sc20012-il)
Reynolds number: 50,000
Max Cl/Cd: 24.55 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20012-il-50000.txt
Download as CSV file: xf-sc20012-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0012 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5509   0.11345   0.10572   0.0195   1.0000   0.3818
  -9.250  -0.5467   0.11006   0.10236   0.0198   1.0000   0.3968
  -9.000  -0.8009   0.07272   0.06476  -0.0197   1.0000   0.1676
  -8.750  -0.7924   0.06782   0.05978  -0.0195   1.0000   0.1650
  -8.500  -0.7935   0.06290   0.05465  -0.0188   1.0000   0.1621
  -8.250  -0.7968   0.05822   0.04964  -0.0176   1.0000   0.1606
  -8.000  -0.7953   0.05408   0.04511  -0.0162   1.0000   0.1610
  -7.750  -0.7897   0.05028   0.04089  -0.0146   1.0000   0.1620
  -7.500  -0.7806   0.04676   0.03690  -0.0130   1.0000   0.1634
  -7.250  -0.7706   0.04372   0.03325  -0.0111   1.0000   0.1665
  -7.000  -0.7510   0.04096   0.03046  -0.0102   1.0000   0.1713
  -6.750  -0.7324   0.03864   0.02789  -0.0090   1.0000   0.1769
  -6.500  -0.7147   0.03637   0.02515  -0.0075   1.0000   0.1830
  -6.250  -0.6920   0.03437   0.02318  -0.0066   1.0000   0.1903
  -6.000  -0.6714   0.03254   0.02109  -0.0054   1.0000   0.1990
  -5.750  -0.6486   0.03090   0.01942  -0.0044   1.0000   0.2089
  -5.500  -0.6256   0.02934   0.01787  -0.0034   1.0000   0.2211
  -5.250  -0.6015   0.02789   0.01647  -0.0024   1.0000   0.2357
  -5.000  -0.5778   0.02657   0.01525  -0.0014   1.0000   0.2560
  -4.750  -0.5544   0.02520   0.01413  -0.0002   1.0000   0.2840
  -4.500  -0.5357   0.02369   0.01304   0.0016   1.0000   0.3298
  -4.250  -0.5274   0.02149   0.01202   0.0052   1.0000   0.4366
  -4.000  -0.5289   0.02205   0.01415   0.0165   1.0000   0.7110
  -3.750  -0.5034   0.02437   0.01636   0.0242   1.0000   0.8006
  -3.500  -0.3159   0.02877   0.01975   0.0077   1.0000   0.8980
  -3.250  -0.1718   0.02855   0.01892  -0.0111   1.0000   0.9591
  -3.000  -0.0646   0.02676   0.01679  -0.0273   1.0000   1.0000
  -2.750  -0.0486   0.02611   0.01607  -0.0271   1.0000   1.0000
  -2.500  -0.0332   0.02556   0.01548  -0.0267   1.0000   1.0000
  -2.250  -0.0190   0.02511   0.01501  -0.0260   1.0000   1.0000
  -2.000  -0.0067   0.02478   0.01467  -0.0249   1.0000   1.0000
  -1.750   0.0020   0.02459   0.01447  -0.0232   1.0000   1.0000
  -1.500   0.0060   0.02453   0.01442  -0.0206   1.0000   1.0000
  -1.250   0.0068   0.02455   0.01444  -0.0175   1.0000   1.0000
  -1.000   0.0058   0.02462   0.01450  -0.0141   1.0000   1.0000
  -0.750   0.0041   0.02470   0.01457  -0.0105   1.0000   1.0000
  -0.500   0.0026   0.02476   0.01461  -0.0070   1.0000   1.0000
  -0.250   0.0012   0.02479   0.01464  -0.0035   1.0000   1.0000
   0.000   0.0000   0.02480   0.01465   0.0000   1.0000   1.0000
   0.250  -0.0012   0.02479   0.01464   0.0035   1.0000   1.0000
   0.500  -0.0026   0.02475   0.01461   0.0070   1.0000   1.0000
   0.750  -0.0041   0.02469   0.01456   0.0105   1.0000   1.0000
   1.000  -0.0058   0.02462   0.01450   0.0141   1.0000   1.0000
   1.250  -0.0068   0.02455   0.01443   0.0175   1.0000   1.0000
   1.500  -0.0060   0.02452   0.01441   0.0206   1.0000   1.0000
   1.750  -0.0020   0.02458   0.01446   0.0232   1.0000   1.0000
   2.000   0.0068   0.02477   0.01466   0.0249   1.0000   1.0000
   2.250   0.0190   0.02510   0.01500   0.0260   1.0000   1.0000
   2.500   0.0334   0.02555   0.01547   0.0267   1.0000   1.0000
   2.750   0.0488   0.02609   0.01606   0.0271   1.0000   1.0000
   3.000   0.0648   0.02675   0.01677   0.0272   1.0000   1.0000
   3.250   0.1719   0.02853   0.01890   0.0111   0.9591   1.0000
   3.500   0.3160   0.02876   0.01974  -0.0077   0.8980   1.0000
   3.750   0.5033   0.02437   0.01635  -0.0242   0.8007   1.0000
   4.000   0.5288   0.02204   0.01414  -0.0165   0.7112   1.0000
   4.250   0.5274   0.02148   0.01202  -0.0051   0.4370   1.0000
   4.500   0.5356   0.02369   0.01304  -0.0015   0.3299   1.0000
   4.750   0.5544   0.02519   0.01412   0.0002   0.2841   1.0000
   5.000   0.5777   0.02657   0.01525   0.0014   0.2560   1.0000
   5.250   0.6015   0.02789   0.01646   0.0025   0.2357   1.0000
   5.500   0.6256   0.02934   0.01787   0.0034   0.2211   1.0000
   5.750   0.6485   0.03090   0.01941   0.0044   0.2089   1.0000
   6.000   0.6714   0.03254   0.02108   0.0054   0.1990   1.0000
   6.250   0.6920   0.03437   0.02318   0.0067   0.1903   1.0000
   6.500   0.7147   0.03637   0.02515   0.0075   0.1830   1.0000
   6.750   0.7324   0.03864   0.02789   0.0090   0.1769   1.0000
   7.000   0.7510   0.04096   0.03046   0.0102   0.1713   1.0000
   7.250   0.7706   0.04372   0.03325   0.0111   0.1665   1.0000
   7.500   0.7807   0.04676   0.03690   0.0130   0.1634   1.0000
   7.750   0.7897   0.05028   0.04089   0.0146   0.1620   1.0000
   8.000   0.7954   0.05408   0.04511   0.0161   0.1610   1.0000
   8.250   0.7969   0.05822   0.04964   0.0176   0.1606   1.0000
   8.500   0.7936   0.06290   0.05466   0.0187   0.1621   1.0000
   8.750   0.7926   0.06784   0.05979   0.0194   0.1650   1.0000
   9.000   0.8029   0.07276   0.06475   0.0197   0.1672   1.0000
   9.250   0.7013   0.08478   0.07727   0.0153   0.1937   1.0000
   9.500   0.5521   0.11348   0.10575  -0.0196   0.3817   1.0000
<< Back to NASA SC(2)-0012 AIRFOIL (sc20012-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-0012 AIRFOIL (sc20012-il)