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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
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