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NACA-0009 9.0% smoothed (n0009sm-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA-0009 9.0% smoothed (n0009sm-il)
Reynolds number: 100,000
Max Cl/Cd: 37.3 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n0009sm-il-100000.txt
Download as CSV file: xf-n0009sm-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA-0009 9.0% smoothed                         
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.6402   0.09783   0.09278  -0.0047   1.0000   0.1611
  -9.500  -0.7228   0.07389   0.06897  -0.0257   1.0000   0.0911
  -9.250  -0.7280   0.07174   0.06685  -0.0244   1.0000   0.0992
  -9.000  -0.8160   0.05631   0.05037  -0.0241   1.0000   0.0670
  -8.750  -0.8199   0.05152   0.04526  -0.0220   1.0000   0.0667
  -8.500  -0.8242   0.04803   0.04118  -0.0189   1.0000   0.0677
  -8.250  -0.8075   0.04520   0.03868  -0.0187   1.0000   0.0739
  -8.000  -0.8037   0.04147   0.03449  -0.0160   1.0000   0.0757
  -7.750  -0.7970   0.03798   0.03045  -0.0132   1.0000   0.0777
  -7.500  -0.7890   0.03532   0.02710  -0.0102   1.0000   0.0816
  -7.250  -0.7711   0.03273   0.02452  -0.0090   1.0000   0.0861
  -7.000  -0.7540   0.03062   0.02204  -0.0071   1.0000   0.0904
  -6.750  -0.7372   0.02863   0.01959  -0.0051   1.0000   0.0958
  -6.500  -0.7158   0.02676   0.01761  -0.0039   1.0000   0.1004
  -6.250  -0.6939   0.02517   0.01574  -0.0025   1.0000   0.1053
  -6.000  -0.6719   0.02363   0.01415  -0.0014   1.0000   0.1123
  -5.750  -0.6496   0.02250   0.01283  -0.0001   1.0000   0.1202
  -5.500  -0.6265   0.02104   0.01145   0.0008   1.0000   0.1273
  -5.250  -0.6042   0.01990   0.01027   0.0021   1.0000   0.1373
  -5.000  -0.5836   0.01889   0.00933   0.0035   1.0000   0.1519
  -4.750  -0.5639   0.01787   0.00846   0.0051   1.0000   0.1707
  -4.500  -0.5459   0.01676   0.00758   0.0071   1.0000   0.2011
  -4.250  -0.5316   0.01542   0.00682   0.0095   1.0000   0.2743
  -4.000  -0.5198   0.01424   0.00639   0.0125   1.0000   0.3983
  -3.750  -0.5050   0.01356   0.00614   0.0152   1.0000   0.4932
  -3.500  -0.4884   0.01309   0.00595   0.0179   1.0000   0.5720
  -3.250  -0.4707   0.01273   0.00586   0.0206   1.0000   0.6441
  -3.000  -0.4511   0.01247   0.00584   0.0232   1.0000   0.7119
  -2.750  -0.4276   0.01237   0.00591   0.0252   1.0000   0.7777
  -2.500  -0.3961   0.01253   0.00614   0.0261   1.0000   0.8439
  -2.250  -0.3457   0.01306   0.00661   0.0235   1.0000   0.8996
  -2.000  -0.2781   0.01364   0.00698   0.0169   1.0000   0.9314
  -1.750  -0.2084   0.01399   0.00711   0.0092   1.0000   0.9547
  -1.500  -0.1341   0.01406   0.00699   0.0001   1.0000   0.9730
  -1.250  -0.0629   0.01388   0.00666  -0.0089   1.0000   0.9911
  -1.000  -0.0197   0.01352   0.00623  -0.0132   1.0000   1.0000
  -0.750  -0.0138   0.01323   0.00594  -0.0103   1.0000   1.0000
  -0.500  -0.0085   0.01303   0.00575  -0.0071   1.0000   1.0000
  -0.250  -0.0040   0.01291   0.00563  -0.0036   1.0000   1.0000
   0.000   0.0000   0.01287   0.00559   0.0000   1.0000   1.0000
   0.250   0.0040   0.01291   0.00563   0.0036   1.0000   1.0000
   0.500   0.0085   0.01303   0.00575   0.0071   1.0000   1.0000
   0.750   0.0138   0.01323   0.00594   0.0103   1.0000   1.0000
   1.000   0.0197   0.01352   0.00623   0.0132   1.0000   1.0000
   1.250   0.0628   0.01387   0.00665   0.0090   0.9911   1.0000
   1.500   0.1342   0.01406   0.00699  -0.0001   0.9730   1.0000
   1.750   0.2086   0.01399   0.00710  -0.0093   0.9546   1.0000
   2.000   0.2783   0.01363   0.00698  -0.0170   0.9314   1.0000
   2.250   0.3459   0.01306   0.00661  -0.0235   0.8995   1.0000
   2.500   0.3960   0.01253   0.00614  -0.0261   0.8439   1.0000
   2.750   0.4275   0.01237   0.00591  -0.0252   0.7775   1.0000
   3.000   0.4510   0.01247   0.00583  -0.0231   0.7118   1.0000
   3.250   0.4705   0.01273   0.00586  -0.0206   0.6438   1.0000
   3.500   0.4883   0.01309   0.00595  -0.0179   0.5719   1.0000
   3.750   0.5049   0.01356   0.00614  -0.0152   0.4935   1.0000
   4.000   0.5197   0.01424   0.00638  -0.0124   0.3979   1.0000
   4.250   0.5315   0.01541   0.00681  -0.0095   0.2749   1.0000
   4.500   0.5458   0.01676   0.00758  -0.0070   0.2010   1.0000
   4.750   0.5638   0.01787   0.00846  -0.0051   0.1704   1.0000
   5.000   0.5835   0.01890   0.00933  -0.0035   0.1518   1.0000
   5.250   0.6041   0.01990   0.01026  -0.0021   0.1372   1.0000
   5.500   0.6264   0.02104   0.01145  -0.0008   0.1274   1.0000
   5.750   0.6495   0.02250   0.01282   0.0002   0.1202   1.0000
   6.000   0.6719   0.02364   0.01416   0.0015   0.1124   1.0000
   6.250   0.6938   0.02517   0.01574   0.0025   0.1054   1.0000
   6.500   0.7158   0.02676   0.01761   0.0040   0.1004   1.0000
   6.750   0.7371   0.02863   0.01959   0.0051   0.0958   1.0000
   7.000   0.7540   0.03063   0.02205   0.0071   0.0905   1.0000
   7.250   0.7711   0.03274   0.02453   0.0091   0.0862   1.0000
   7.500   0.7891   0.03528   0.02705   0.0102   0.0817   1.0000
   7.750   0.7971   0.03798   0.03044   0.0132   0.0777   1.0000
   8.000   0.8036   0.04149   0.03451   0.0160   0.0756   1.0000
   8.250   0.8073   0.04527   0.03876   0.0187   0.0741   1.0000
   8.500   0.8242   0.04805   0.04120   0.0189   0.0677   1.0000
   8.750   0.8197   0.05156   0.04531   0.0220   0.0667   1.0000
   9.000   0.8155   0.05629   0.05037   0.0242   0.0669   1.0000
   9.250   0.8169   0.06167   0.05590   0.0254   0.0680   1.0000
   9.500   0.6962   0.09159   0.08675   0.0152   0.1627   1.0000
   9.750   0.6376   0.09795   0.09289   0.0041   0.1607   1.0000
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