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NACA 16-012 (naca16012-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 16-012 (naca16012-il)
Reynolds number: 50,000
Max Cl/Cd: 21.08 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca16012-il-50000.txt
Download as CSV file: xf-naca16012-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 16-012                                     
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5757   0.07902   0.07264  -0.0301   1.0000   0.2245
  -8.750  -0.6281   0.07381   0.06748  -0.0284   1.0000   0.2159
  -8.500  -0.6730   0.07109   0.06482  -0.0226   1.0000   0.2174
  -8.250  -0.7236   0.06867   0.06238  -0.0158   1.0000   0.2190
  -8.000  -0.7522   0.06081   0.05416  -0.0131   1.0000   0.1734
  -7.750  -0.8322   0.06565   0.05805  -0.0042   1.0000   0.1567
  -7.500  -0.8435   0.06154   0.05342   0.0009   1.0000   0.1440
  -7.250  -0.8426   0.05763   0.04922   0.0047   1.0000   0.1390
  -7.000  -0.8471   0.05432   0.04516   0.0102   1.0000   0.1311
  -6.750  -0.8376   0.05062   0.04119   0.0131   1.0000   0.1281
  -6.500  -0.8287   0.04727   0.03735   0.0167   1.0000   0.1249
  -6.250  -0.8166   0.04429   0.03387   0.0198   1.0000   0.1241
  -6.000  -0.8020   0.04183   0.03099   0.0225   1.0000   0.1276
  -5.750  -0.7838   0.03943   0.02806   0.0249   1.0000   0.1303
  -5.500  -0.7609   0.03714   0.02521   0.0267   1.0000   0.1324
  -5.250  -0.7328   0.03476   0.02279   0.0267   1.0000   0.1415
  -5.000  -0.2010   0.03446   0.02419  -0.0441   1.0000   1.0000
  -4.750  -0.1928   0.03365   0.02318  -0.0425   1.0000   1.0000
  -4.500  -0.1841   0.03294   0.02226  -0.0408   1.0000   1.0000
  -4.250  -0.1750   0.03231   0.02146  -0.0389   1.0000   1.0000
  -4.000  -0.1656   0.03176   0.02075  -0.0370   1.0000   1.0000
  -3.750  -0.1559   0.03127   0.02011  -0.0350   1.0000   1.0000
  -3.500  -0.1460   0.03083   0.01954  -0.0329   1.0000   1.0000
  -3.250  -0.1360   0.03044   0.01903  -0.0308   1.0000   1.0000
  -3.000  -0.1258   0.03009   0.01855  -0.0286   1.0000   1.0000
  -2.750  -0.1156   0.02979   0.01814  -0.0264   1.0000   1.0000
  -2.500  -0.1052   0.02952   0.01778  -0.0241   1.0000   1.0000
  -2.250  -0.0949   0.02928   0.01745  -0.0218   1.0000   1.0000
  -2.000  -0.0844   0.02907   0.01717  -0.0194   1.0000   1.0000
  -1.750  -0.0739   0.02889   0.01691  -0.0171   1.0000   1.0000
  -1.500  -0.0634   0.02874   0.01670  -0.0147   1.0000   1.0000
  -1.250  -0.0529   0.02861   0.01653  -0.0122   1.0000   1.0000
  -1.000  -0.0423   0.02851   0.01639  -0.0098   1.0000   1.0000
  -0.750  -0.0318   0.02843   0.01628  -0.0074   1.0000   1.0000
  -0.500  -0.0212   0.02837   0.01620  -0.0049   1.0000   1.0000
  -0.250  -0.0106   0.02834   0.01615  -0.0025   1.0000   1.0000
   0.000   0.0000   0.02833   0.01613   0.0000   1.0000   1.0000
   0.250   0.0106   0.02834   0.01615   0.0025   1.0000   1.0000
   0.500   0.0212   0.02837   0.01620   0.0049   1.0000   1.0000
   0.750   0.0318   0.02842   0.01628   0.0074   1.0000   1.0000
   1.000   0.0423   0.02850   0.01638   0.0098   1.0000   1.0000
   1.250   0.0529   0.02860   0.01652   0.0122   1.0000   1.0000
   1.500   0.0634   0.02873   0.01669   0.0147   1.0000   1.0000
   1.750   0.0739   0.02888   0.01690   0.0171   1.0000   1.0000
   2.000   0.0844   0.02906   0.01716   0.0194   1.0000   1.0000
   2.250   0.0949   0.02926   0.01744   0.0218   1.0000   1.0000
   2.500   0.1053   0.02950   0.01776   0.0241   1.0000   1.0000
   2.750   0.1156   0.02977   0.01812   0.0264   1.0000   1.0000
   3.000   0.1258   0.03007   0.01853   0.0286   1.0000   1.0000
   3.250   0.1360   0.03042   0.01901   0.0308   1.0000   1.0000
   3.500   0.1460   0.03080   0.01952   0.0329   1.0000   1.0000
   3.750   0.1559   0.03124   0.02009   0.0350   1.0000   1.0000
   4.000   0.1656   0.03173   0.02072   0.0370   1.0000   1.0000
   4.250   0.1751   0.03228   0.02143   0.0390   1.0000   1.0000
   4.500   0.1842   0.03290   0.02222   0.0408   1.0000   1.0000
   4.750   0.1929   0.03361   0.02314   0.0425   1.0000   1.0000
   5.000   0.2011   0.03442   0.02415   0.0441   1.0000   1.0000
   5.250   0.7326   0.03475   0.02278  -0.0267   0.1416   1.0000
   5.500   0.7607   0.03713   0.02519  -0.0267   0.1324   1.0000
   5.750   0.7836   0.03942   0.02804  -0.0249   0.1304   1.0000
   6.000   0.8019   0.04182   0.03098  -0.0225   0.1276   1.0000
   6.250   0.8165   0.04427   0.03386  -0.0198   0.1242   1.0000
   6.500   0.8286   0.04725   0.03733  -0.0166   0.1249   1.0000
   6.750   0.8375   0.05061   0.04117  -0.0131   0.1281   1.0000
   7.000   0.8470   0.05430   0.04514  -0.0102   0.1311   1.0000
   7.250   0.8425   0.05761   0.04921  -0.0046   0.1389   1.0000
   7.500   0.8433   0.06152   0.05340  -0.0009   0.1439   1.0000
   7.750   0.7731   0.05584   0.04888   0.0083   0.1587   1.0000
   8.250   0.7450   0.08115   0.07465   0.0094   0.2910   1.0000
   8.500   0.6960   0.08351   0.07690   0.0152   0.2886   1.0000
   8.750   0.6270   0.08900   0.08215   0.0154   0.3151   1.0000
   9.250   0.5340   0.10363   0.09637   0.0045   0.4403   1.0000
   9.500   0.5338   0.10681   0.09950   0.0063   0.4221   1.0000
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