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NACA 0010-34 a=0.8 c(li)=0.2 (naca001034a08cli02-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA 0010-34 a=0.8 c(li)=0.2 (naca001034a08cli02-il)
Reynolds number: 500,000
Max Cl/Cd: 89.59 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca001034a08cli02-il-500000.txt
Download as CSV file: xf-naca001034a08cli02-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0010-34 a=0.8 c(li)=0.2                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5057   0.08467   0.08250  -0.0416   1.0000   0.0101
  -9.250  -0.5107   0.07880   0.07666  -0.0457   1.0000   0.0102
  -9.000  -0.5174   0.07277   0.07062  -0.0510   1.0000   0.0100
  -8.750  -0.5338   0.06986   0.06769  -0.0504   1.0000   0.0099
  -8.500  -0.5520   0.06915   0.06697  -0.0459   1.0000   0.0098
  -8.250  -0.5741   0.06774   0.06554  -0.0408   1.0000   0.0098
  -8.000  -0.5936   0.06518   0.06292  -0.0369   1.0000   0.0098
  -7.750  -0.6000   0.05819   0.05573  -0.0390   0.9973   0.0101
  -7.500  -0.5931   0.05237   0.04970  -0.0417   0.9944   0.0105
  -7.250  -0.5802   0.04855   0.04572  -0.0431   0.9912   0.0110
  -7.000  -0.5624   0.04500   0.04198  -0.0447   0.9887   0.0117
  -6.750  -0.5411   0.04144   0.03819  -0.0463   0.9869   0.0126
  -6.500  -0.5232   0.03839   0.03490  -0.0461   0.9829   0.0141
  -6.250  -0.4879   0.03841   0.03466  -0.0464   0.9809   0.0180
  -6.000  -0.4612   0.03600   0.03187  -0.0470   0.9789   0.0183
  -5.750  -0.4345   0.03290   0.02842  -0.0478   0.9775   0.0184
  -5.500  -0.4269   0.02591   0.02087  -0.0456   0.9720   0.0198
  -5.250  -0.4000   0.02327   0.01803  -0.0463   0.9700   0.0212
  -5.000  -0.3691   0.02119   0.01573  -0.0471   0.9686   0.0224
  -4.750  -0.3353   0.01697   0.01102  -0.0462   0.9680   0.0127
  -4.500  -0.3022   0.01509   0.00895  -0.0470   0.9672   0.0127
  -4.250  -0.2761   0.01407   0.00784  -0.0467   0.9640   0.0139
  -4.000  -0.2478   0.01334   0.00706  -0.0469   0.9608   0.0160
  -3.750  -0.2169   0.01238   0.00599  -0.0476   0.9588   0.0173
  -3.500  -0.1853   0.01157   0.00504  -0.0485   0.9570   0.0193
  -3.250  -0.1531   0.01084   0.00421  -0.0496   0.9554   0.0280
  -3.000  -0.1336   0.00939   0.00356  -0.0485   0.9509   0.2417
  -2.750  -0.1159   0.00801   0.00330  -0.0472   0.9461   0.5329
  -2.500  -0.0902   0.00743   0.00317  -0.0469   0.9433   0.6650
  -2.250  -0.0630   0.00709   0.00310  -0.0465   0.9410   0.7537
  -2.000  -0.0418   0.00696   0.00309  -0.0448   0.9350   0.8031
  -1.750  -0.0148   0.00681   0.00301  -0.0444   0.9313   0.8373
  -1.500   0.0147   0.00669   0.00289  -0.0445   0.9285   0.8563
  -1.250   0.0398   0.00664   0.00284  -0.0438   0.9228   0.8743
  -1.000   0.0672   0.00656   0.00278  -0.0434   0.9182   0.8918
  -0.750   0.0965   0.00650   0.00270  -0.0435   0.9148   0.9073
  -0.500   0.1223   0.00650   0.00270  -0.0429   0.9084   0.9218
  -0.250   0.1516   0.00648   0.00266  -0.0431   0.9038   0.9334
   0.000   0.1817   0.00647   0.00264  -0.0434   0.8993   0.9438
   0.250   0.2101   0.00649   0.00265  -0.0435   0.8928   0.9543
   0.500   0.2452   0.00648   0.00262  -0.0450   0.8884   0.9600
   0.750   0.2770   0.00650   0.00265  -0.0459   0.8818   0.9676
   1.000   0.3148   0.00649   0.00263  -0.0481   0.8761   0.9715
   1.250   0.3496   0.00650   0.00266  -0.0497   0.8689   0.9775
   1.500   0.3893   0.00647   0.00263  -0.0523   0.8617   0.9805
   1.750   0.4253   0.00647   0.00265  -0.0542   0.8512   0.9860
   2.000   0.4638   0.00644   0.00266  -0.0567   0.8405   0.9893
   2.250   0.5007   0.00641   0.00263  -0.0587   0.8241   0.9936
   2.500   0.5367   0.00638   0.00253  -0.0605   0.7905   0.9972
   2.750   0.5708   0.00644   0.00252  -0.0619   0.7533   1.0000
   3.000   0.5922   0.00661   0.00251  -0.0607   0.7015   1.0000
   3.250   0.6114   0.00697   0.00257  -0.0590   0.6289   1.0000
   3.500   0.6270   0.00758   0.00275  -0.0567   0.5261   1.0000
   3.750   0.6383   0.00854   0.00307  -0.0538   0.3852   1.0000
   4.000   0.6425   0.01024   0.00366  -0.0500   0.1532   1.0000
   4.250   0.6526   0.01149   0.00435  -0.0470   0.0341   1.0000
   4.500   0.6706   0.01204   0.00496  -0.0450   0.0236   1.0000
   4.750   0.6871   0.01273   0.00570  -0.0427   0.0176   1.0000
   5.000   0.6995   0.01381   0.00692  -0.0396   0.0153   1.0000
   5.250   0.7164   0.01451   0.00771  -0.0373   0.0142   1.0000
   5.500   0.7338   0.01522   0.00849  -0.0353   0.0125   1.0000
   5.750   0.7516   0.01597   0.00929  -0.0334   0.0110   1.0000
   6.000   0.7692   0.01703   0.01043  -0.0313   0.0102   1.0000
   6.250   0.7880   0.01840   0.01189  -0.0296   0.0095   1.0000
   6.500   0.8097   0.02034   0.01404  -0.0281   0.0094   1.0000
   6.750   0.8364   0.02582   0.02008  -0.0261   0.0127   1.0000
   7.500   0.8607   0.04126   0.03660  -0.0165   0.0157   1.0000
   7.750   0.8676   0.04423   0.03985  -0.0135   0.0157   1.0000
   8.000   0.8728   0.04700   0.04290  -0.0104   0.0156   1.0000
   8.250   0.8768   0.04950   0.04567  -0.0072   0.0155   1.0000
   8.500   0.9032   0.04814   0.04452  -0.0054   0.0131   1.0000
   8.750   0.9055   0.05122   0.04784  -0.0023   0.0120   1.0000
   9.000   0.9033   0.05445   0.05129   0.0007   0.0114   1.0000
   9.250   0.8972   0.05769   0.05471   0.0037   0.0109   1.0000
   9.500   0.8860   0.06075   0.05792   0.0071   0.0106   1.0000
   9.750   0.8703   0.06385   0.06116   0.0104   0.0105   1.0000
  10.000   0.8527   0.06739   0.06484   0.0125   0.0104   1.0000
  10.250   0.8332   0.07164   0.06921   0.0131   0.0104   1.0000
  10.500   0.8121   0.07694   0.07464   0.0119   0.0106   1.0000
  10.750   0.7898   0.08394   0.08174   0.0079   0.0109   1.0000
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