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NACA 65(1)-412 (naca651412-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 65(1)-412 (naca651412-il)
Reynolds number: 100,000
Max Cl/Cd: 52.98 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca651412-il-100000.txt
Download as CSV file: xf-naca651412-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 65(1)-412                                  
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.4187   0.09531   0.09040  -0.0462   1.0000   0.1221
  -9.250  -0.4491   0.09178   0.08705  -0.0503   1.0000   0.1257
  -9.000  -0.4840   0.08843   0.08386  -0.0527   1.0000   0.1261
  -8.750  -0.4399   0.08610   0.08147  -0.0463   1.0000   0.1364
  -8.500  -0.4676   0.08335   0.07887  -0.0462   1.0000   0.1385
  -8.250  -0.4966   0.08129   0.07695  -0.0437   1.0000   0.1383
  -8.000  -0.5279   0.07959   0.07536  -0.0404   1.0000   0.1379
  -7.750  -0.5599   0.07751   0.07332  -0.0382   1.0000   0.1387
  -7.500  -0.5999   0.07559   0.07132  -0.0370   1.0000   0.1404
  -7.250  -0.6076   0.07194   0.06772  -0.0350   1.0000   0.1437
  -7.000  -0.6006   0.07009   0.06594  -0.0319   1.0000   0.1503
  -6.750  -0.6271   0.06678   0.06239  -0.0324   1.0000   0.1577
  -6.500  -0.6164   0.06476   0.06051  -0.0293   1.0000   0.1666
  -6.000  -0.5851   0.05792   0.05353  -0.0331   0.9913   0.2074
  -5.250  -0.4736   0.03763   0.02968  -0.0465   0.9751   0.0832
  -5.000  -0.4391   0.03427   0.02599  -0.0486   0.9716   0.0821
  -4.750  -0.4081   0.03165   0.02306  -0.0496   0.9674   0.0799
  -4.500  -0.3762   0.02967   0.02074  -0.0503   0.9625   0.0785
  -4.250  -0.3393   0.02813   0.01887  -0.0519   0.9584   0.0795
  -4.000  -0.3051   0.02726   0.01771  -0.0531   0.9539   0.0829
  -3.750  -0.2781   0.02593   0.01635  -0.0530   0.9488   0.0850
  -3.500  -0.2451   0.02482   0.01536  -0.0542   0.9444   0.0892
  -3.250  -0.2072   0.02414   0.01464  -0.0563   0.9410   0.0985
  -3.000  -0.1882   0.02344   0.01407  -0.0552   0.9341   0.1072
  -2.750  -0.1559   0.02275   0.01341  -0.0565   0.9296   0.1275
  -2.500  -0.1366   0.02042   0.01382  -0.0546   0.9265   0.6949
  -2.250  -0.1330   0.02116   0.01463  -0.0482   0.9186   0.7586
  -2.000  -0.1227   0.02202   0.01552  -0.0421   0.9130   0.8079
  -1.750  -0.1224   0.02262   0.01612  -0.0351   0.9063   0.8436
  -1.500  -0.1226   0.02309   0.01660  -0.0273   0.8999   0.8802
  -1.250  -0.1010   0.02364   0.01708  -0.0227   0.8964   0.9258
  -1.000  -0.0720   0.02383   0.01713  -0.0232   0.8910   0.9464
  -0.750  -0.0241   0.02398   0.01710  -0.0280   0.8869   0.9526
  -0.500   0.0221   0.02410   0.01706  -0.0324   0.8833   0.9599
  -0.250   0.0814   0.02428   0.01710  -0.0393   0.8813   0.9638
   0.000   0.1123   0.02450   0.01726  -0.0416   0.8748   0.9727
   0.250   0.1657   0.02467   0.01734  -0.0476   0.8711   0.9775
   0.500   0.2180   0.02480   0.01741  -0.0533   0.8681   0.9832
   0.750   0.2713   0.02495   0.01752  -0.0592   0.8651   0.9881
   1.000   0.3053   0.02529   0.01786  -0.0623   0.8586   0.9980
   1.250   0.3351   0.02544   0.01801  -0.0639   0.8537   1.0000
   1.500   0.3140   0.02584   0.01840  -0.0571   0.8439   1.0000
   1.750   0.3315   0.02601   0.01856  -0.0563   0.8384   1.0000
   2.000   0.3191   0.02644   0.01897  -0.0511   0.8291   1.0000
   2.250   0.3504   0.02672   0.01926  -0.0527   0.8240   1.0000
   2.500   0.3707   0.02724   0.01978  -0.0527   0.8170   1.0000
   2.750   0.3996   0.02764   0.02022  -0.0539   0.8104   1.0000
   3.000   0.4452   0.02779   0.02042  -0.0572   0.8068   1.0000
   3.250   0.4548   0.02861   0.02126  -0.0557   0.7965   1.0000
   3.500   0.4992   0.02872   0.02147  -0.0586   0.7925   1.0000
   3.750   0.5118   0.02957   0.02235  -0.0575   0.7821   1.0000
   4.000   0.5605   0.02941   0.02232  -0.0605   0.7775   1.0000
   4.250   0.5802   0.02989   0.02289  -0.0598   0.7657   1.0000
   4.500   0.6110   0.03003   0.02317  -0.0603   0.7558   1.0000
   4.750   0.6605   0.02940   0.02274  -0.0628   0.7493   1.0000
   5.000   0.7603   0.02391   0.01760  -0.0667   0.7304   1.0000
   5.250   0.8277   0.01875   0.01252  -0.0657   0.6899   1.0000
   5.500   0.8562   0.01748   0.01137  -0.0636   0.6553   1.0000
   5.750   0.8720   0.01646   0.01029  -0.0592   0.5776   1.0000
   6.000   0.8524   0.01827   0.01015  -0.0502   0.2672   1.0000
   6.250   0.8386   0.02126   0.01188  -0.0446   0.1258   1.0000
   6.500   0.8446   0.02284   0.01324  -0.0415   0.1022   1.0000
   6.750   0.8531   0.02430   0.01460  -0.0388   0.0915   1.0000
   7.000   0.8680   0.02559   0.01585  -0.0370   0.0825   1.0000
   7.250   0.8902   0.02718   0.01740  -0.0363   0.0760   1.0000
   7.500   0.9240   0.02890   0.01910  -0.0371   0.0711   1.0000
   7.750   0.9717   0.03191   0.02194  -0.0402   0.0664   1.0000
   8.000   1.0008   0.03352   0.02385  -0.0402   0.0635   1.0000
   8.250   1.0343   0.03603   0.02662  -0.0408   0.0623   1.0000
   8.500   1.0632   0.03889   0.02983  -0.0408   0.0621   1.0000
   8.750   1.0875   0.04212   0.03345  -0.0401   0.0629   1.0000
   9.000   1.1065   0.04551   0.03724  -0.0389   0.0635   1.0000
   9.250   1.1200   0.04883   0.04099  -0.0372   0.0636   1.0000
   9.500   1.1305   0.05251   0.04504  -0.0354   0.0642   1.0000
   9.750   1.1533   0.05894   0.05156  -0.0358   0.0672   1.0000
  10.000   1.0904   0.05766   0.05151  -0.0267   0.0799   1.0000
  12.000   0.8292   0.13212   0.12757  -0.0484   0.1680   1.0000
  12.250   0.6746   0.13382   0.12968  -0.0425   0.1670   1.0000
  12.500   0.6939   0.13715   0.13304  -0.0416   0.1636   1.0000
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