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

NACA-0009 9.0% smoothed (n0009sm-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA-0009 9.0% smoothed (n0009sm-il)
Reynolds number: 50,000
Max Cl/Cd: 27.03 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n0009sm-il-50000.txt
Download as CSV file: xf-n0009sm-il-50000.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 =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5693   0.10116   0.09376   0.0165   1.0000   0.3639
  -8.750  -0.7142   0.07868   0.07178  -0.0149   1.0000   0.1757
  -8.500  -0.7137   0.07346   0.06654  -0.0151   1.0000   0.1724
  -8.250  -0.7642   0.06347   0.05612  -0.0184   1.0000   0.1482
  -8.000  -0.7681   0.05866   0.05100  -0.0173   1.0000   0.1481
  -7.750  -0.7702   0.05414   0.04605  -0.0158   1.0000   0.1486
  -7.500  -0.7678   0.04985   0.04126  -0.0139   1.0000   0.1489
  -7.250  -0.7586   0.04570   0.03671  -0.0122   1.0000   0.1494
  -7.000  -0.7428   0.04225   0.03304  -0.0107   1.0000   0.1522
  -6.750  -0.7290   0.03964   0.03006  -0.0089   1.0000   0.1595
  -6.500  -0.7134   0.03689   0.02699  -0.0071   1.0000   0.1662
  -6.250  -0.6942   0.03431   0.02407  -0.0056   1.0000   0.1718
  -6.000  -0.6744   0.03181   0.02110  -0.0040   1.0000   0.1782
  -5.750  -0.6537   0.03001   0.01911  -0.0026   1.0000   0.1915
  -5.500  -0.6311   0.02822   0.01724  -0.0014   1.0000   0.2060
  -5.250  -0.6064   0.02638   0.01536  -0.0003   1.0000   0.2220
  -5.000  -0.5831   0.02481   0.01382   0.0009   1.0000   0.2480
  -4.750  -0.5614   0.02330   0.01254   0.0024   1.0000   0.2889
  -4.500  -0.5414   0.02154   0.01133   0.0044   1.0000   0.3597
  -4.250  -0.5291   0.01998   0.01068   0.0084   1.0000   0.4790
  -4.000  -0.5177   0.01915   0.01049   0.0134   1.0000   0.5943
  -3.750  -0.5027   0.01887   0.01051   0.0185   1.0000   0.6872
  -3.500  -0.4803   0.01909   0.01089   0.0231   1.0000   0.7695
  -3.250  -0.4247   0.02039   0.01199   0.0234   1.0000   0.8489
  -3.000  -0.2970   0.02197   0.01277   0.0091   1.0000   0.9159
  -2.750  -0.1843   0.02164   0.01183  -0.0069   1.0000   0.9649
  -2.500  -0.0912   0.02030   0.01010  -0.0216   1.0000   1.0000
  -2.250  -0.0830   0.01956   0.00931  -0.0204   1.0000   1.0000
  -2.000  -0.0746   0.01894   0.00862  -0.0188   1.0000   1.0000
  -1.750  -0.0661   0.01842   0.00807  -0.0170   1.0000   1.0000
  -1.500  -0.0574   0.01800   0.00761  -0.0149   1.0000   1.0000
  -1.250  -0.0485   0.01766   0.00724  -0.0126   1.0000   1.0000
  -1.000  -0.0393   0.01739   0.00695  -0.0102   1.0000   1.0000
  -0.750  -0.0297   0.01719   0.00674  -0.0077   1.0000   1.0000
  -0.500  -0.0199   0.01705   0.00658  -0.0052   1.0000   1.0000
  -0.250  -0.0100   0.01696   0.00649  -0.0026   1.0000   1.0000
   0.000   0.0000   0.01694   0.00646   0.0000   1.0000   1.0000
   0.250   0.0100   0.01696   0.00649   0.0026   1.0000   1.0000
   0.500   0.0199   0.01704   0.00658   0.0052   1.0000   1.0000
   0.750   0.0297   0.01718   0.00674   0.0077   1.0000   1.0000
   1.000   0.0393   0.01739   0.00695   0.0102   1.0000   1.0000
   1.250   0.0485   0.01765   0.00724   0.0126   1.0000   1.0000
   1.500   0.0574   0.01800   0.00760   0.0149   1.0000   1.0000
   1.750   0.0661   0.01842   0.00806   0.0170   1.0000   1.0000
   2.000   0.0747   0.01894   0.00862   0.0188   1.0000   1.0000
   2.250   0.0831   0.01956   0.00931   0.0204   1.0000   1.0000
   2.500   0.0913   0.02030   0.01009   0.0216   1.0000   1.0000
   2.750   0.1843   0.02163   0.01182   0.0070   0.9650   1.0000
   3.000   0.2974   0.02197   0.01276  -0.0091   0.9158   1.0000
   3.250   0.4247   0.02038   0.01198  -0.0234   0.8489   1.0000
   3.500   0.4803   0.01909   0.01088  -0.0231   0.7695   1.0000
   3.750   0.5026   0.01887   0.01052  -0.0185   0.6874   1.0000
   4.000   0.5176   0.01915   0.01049  -0.0134   0.5942   1.0000
   4.250   0.5291   0.01998   0.01068  -0.0083   0.4794   1.0000
   4.500   0.5413   0.02154   0.01133  -0.0044   0.3598   1.0000
   4.750   0.5613   0.02330   0.01253  -0.0024   0.2885   1.0000
   5.000   0.5830   0.02481   0.01382  -0.0009   0.2479   1.0000
   5.250   0.6063   0.02638   0.01535   0.0004   0.2220   1.0000
   5.500   0.6311   0.02822   0.01724   0.0014   0.2060   1.0000
   5.750   0.6536   0.03001   0.01911   0.0026   0.1915   1.0000
   6.000   0.6744   0.03181   0.02109   0.0040   0.1782   1.0000
   6.250   0.6942   0.03431   0.02406   0.0056   0.1718   1.0000
   6.500   0.7133   0.03689   0.02698   0.0072   0.1662   1.0000
   6.750   0.7289   0.03964   0.03006   0.0089   0.1595   1.0000
   7.000   0.7428   0.04226   0.03304   0.0107   0.1522   1.0000
   7.250   0.7590   0.04572   0.03670   0.0121   0.1493   1.0000
   7.500   0.7679   0.04985   0.04125   0.0139   0.1489   1.0000
   7.750   0.7707   0.05416   0.04606   0.0158   0.1486   1.0000
   8.000   0.7677   0.05871   0.05106   0.0174   0.1481   1.0000
   8.250   0.7649   0.06348   0.05611   0.0184   0.1482   1.0000
   8.500   0.7260   0.07196   0.06501   0.0164   0.1658   1.0000
   8.750   0.6992   0.08127   0.07438   0.0121   0.1896   1.0000
   9.000   0.5669   0.10116   0.09375  -0.0169   0.3666   1.0000
<< Back to NACA-0009 9.0% smoothed (n0009sm-il)

Polar data table (+)

Polar graphs


<< Back to NACA-0009 9.0% smoothed (n0009sm-il)