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NACA 64-110 AIRFOIL (n64110-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 64-110 AIRFOIL (n64110-il)
Reynolds number: 100,000
Max Cl/Cd: 41.64 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n64110-il-100000.txt
Download as CSV file: xf-n64110-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 64-110 AIRFOIL                             
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.5973   0.12572   0.12064   0.0047   1.0000   0.0995
 -11.000  -0.5947   0.12194   0.11687   0.0030   1.0000   0.1042
 -10.750  -0.6164   0.11782   0.11287  -0.0040   1.0000   0.1079
 -10.500  -0.6061   0.11271   0.10777  -0.0035   1.0000   0.1104
 -10.250  -0.5907   0.10951   0.10452  -0.0021   1.0000   0.1155
 -10.000  -0.6087   0.10497   0.10010  -0.0079   1.0000   0.1213
  -9.750  -0.6448   0.09885   0.09412  -0.0181   1.0000   0.1222
  -9.500  -0.5931   0.09708   0.09221  -0.0077   1.0000   0.1308
  -9.250  -0.6243   0.09121   0.08648  -0.0164   1.0000   0.1355
  -9.000  -0.5984   0.08870   0.08394  -0.0116   1.0000   0.1436
  -8.750  -0.6199   0.08300   0.07834  -0.0178   1.0000   0.1483
  -8.500  -0.6136   0.07956   0.07494  -0.0170   1.0000   0.1570
  -8.250  -0.6372   0.07446   0.06988  -0.0215   1.0000   0.1614
  -7.750  -0.6451   0.06678   0.06220  -0.0229   1.0000   0.1836
  -6.750  -0.6377   0.04256   0.03594  -0.0282   1.0000   0.1057
  -6.500  -0.6185   0.03617   0.02862  -0.0262   1.0000   0.0765
  -6.250  -0.6001   0.03232   0.02408  -0.0244   1.0000   0.0701
  -6.000  -0.5790   0.03025   0.02145  -0.0226   1.0000   0.0682
  -5.750  -0.5594   0.02774   0.01889  -0.0217   1.0000   0.0712
  -5.500  -0.5386   0.02591   0.01688  -0.0204   1.0000   0.0731
  -5.250  -0.5171   0.02413   0.01487  -0.0189   1.0000   0.0740
  -5.000  -0.4961   0.02263   0.01324  -0.0174   1.0000   0.0760
  -4.750  -0.4761   0.02152   0.01197  -0.0157   1.0000   0.0799
  -4.500  -0.4583   0.02008   0.01067  -0.0141   1.0000   0.0860
  -4.250  -0.4410   0.01910   0.00970  -0.0122   1.0000   0.0921
  -4.000  -0.4257   0.01802   0.00873  -0.0102   1.0000   0.0999
  -3.750  -0.4103   0.01713   0.00791  -0.0084   1.0000   0.1165
  -3.500  -0.3946   0.01590   0.00694  -0.0068   1.0000   0.1548
  -3.250  -0.3946   0.01340   0.00703  -0.0018   1.0000   0.6427
  -3.000  -0.3834   0.01376   0.00745   0.0023   1.0000   0.7092
  -2.750  -0.3718   0.01410   0.00781   0.0062   1.0000   0.7476
  -2.500  -0.3618   0.01445   0.00815   0.0103   1.0000   0.7819
  -2.250  -0.3599   0.01493   0.00869   0.0168   1.0000   0.8209
  -2.000  -0.3456   0.01576   0.00958   0.0222   0.9909   0.8698
  -1.750  -0.3106   0.01628   0.00998   0.0222   0.9829   0.8975
  -1.500  -0.2686   0.01639   0.00993   0.0191   0.9745   0.9109
  -1.250  -0.2180   0.01656   0.00992   0.0142   0.9682   0.9217
  -1.000  -0.1700   0.01663   0.00986   0.0096   0.9606   0.9321
  -0.750  -0.1115   0.01677   0.00988   0.0030   0.9559   0.9406
  -0.500  -0.0588   0.01681   0.00983  -0.0026   0.9488   0.9484
  -0.250  -0.0003   0.01687   0.00982  -0.0093   0.9437   0.9565
   0.000   0.0653   0.01685   0.00978  -0.0174   0.9395   0.9610
   0.250   0.1225   0.01680   0.00973  -0.0240   0.9332   0.9683
   0.500   0.1863   0.01667   0.00962  -0.0316   0.9288   0.9736
   0.750   0.2394   0.01656   0.00957  -0.0374   0.9211   0.9815
   1.000   0.2984   0.01635   0.00944  -0.0442   0.9152   0.9877
   1.250   0.3502   0.01619   0.00938  -0.0497   0.9066   0.9960
   1.500   0.3956   0.01602   0.00934  -0.0537   0.8988   1.0000
   1.750   0.4151   0.01612   0.00952  -0.0530   0.8854   1.0000
   2.000   0.4329   0.01623   0.00970  -0.0517   0.8718   1.0000
   2.250   0.4489   0.01638   0.00992  -0.0501   0.8584   1.0000
   2.500   0.4639   0.01656   0.01019  -0.0481   0.8453   1.0000
   2.750   0.4797   0.01652   0.01025  -0.0455   0.8297   1.0000
   3.000   0.4918   0.01597   0.00972  -0.0406   0.8038   1.0000
   3.250   0.5021   0.01497   0.00866  -0.0341   0.7696   1.0000
   3.500   0.5131   0.01435   0.00798  -0.0289   0.7342   1.0000
   3.750   0.5287   0.01403   0.00769  -0.0254   0.7046   1.0000
   4.000   0.5455   0.01373   0.00742  -0.0221   0.6676   1.0000
   4.250   0.5588   0.01342   0.00693  -0.0177   0.5763   1.0000
   4.500   0.5494   0.01650   0.00746  -0.0114   0.1547   1.0000
   4.750   0.5646   0.01811   0.00870  -0.0096   0.1123   1.0000
   5.000   0.5829   0.01945   0.00990  -0.0081   0.0973   1.0000
   5.250   0.6044   0.02069   0.01108  -0.0070   0.0863   1.0000
   5.500   0.6281   0.02229   0.01262  -0.0062   0.0792   1.0000
   5.750   0.6548   0.02386   0.01424  -0.0056   0.0748   1.0000
   6.000   0.6816   0.02576   0.01612  -0.0053   0.0710   1.0000
   6.250   0.7077   0.02803   0.01854  -0.0049   0.0675   1.0000
   6.500   0.7334   0.03019   0.02105  -0.0041   0.0663   1.0000
   6.750   0.7579   0.03291   0.02416  -0.0033   0.0665   1.0000
   7.000   0.7803   0.03625   0.02784  -0.0024   0.0677   1.0000
   7.250   0.8020   0.04047   0.03230  -0.0019   0.0694   1.0000
   7.500   0.8164   0.04344   0.03636   0.0010   0.0764   1.0000
   7.750   0.8320   0.04848   0.04162   0.0017   0.0810   1.0000
   9.250   0.7557   0.08724   0.08275  -0.0042   0.1638   1.0000
   9.500   0.6159   0.09182   0.08750  -0.0100   0.1780   1.0000
   9.750   0.6451   0.09430   0.09001  -0.0063   0.1652   1.0000
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