Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

OAF117 AIRFOIL (oaf117-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: OAF117 AIRFOIL (oaf117-il)
Reynolds number: 100,000
Max Cl/Cd: 46.95 at α=7.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-oaf117-il-100000.txt
Download as CSV file: xf-oaf117-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: OAF117 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.7083   0.06555   0.06037  -0.0283   1.0000   0.0930
  -8.750  -0.7122   0.05682   0.05132  -0.0349   1.0000   0.0914
  -8.500  -0.7107   0.04819   0.04209  -0.0413   1.0000   0.0913
  -8.250  -0.6978   0.04138   0.03441  -0.0459   1.0000   0.0927
  -8.000  -0.6747   0.03884   0.03194  -0.0464   1.0000   0.0953
  -7.750  -0.6502   0.03610   0.02896  -0.0476   1.0000   0.0980
  -7.500  -0.6242   0.03306   0.02540  -0.0492   1.0000   0.1022
  -7.250  -0.5968   0.03009   0.02192  -0.0507   1.0000   0.1061
  -7.000  -0.5693   0.02859   0.02043  -0.0511   1.0000   0.1106
  -6.750  -0.5399   0.02690   0.01831  -0.0520   1.0000   0.1170
  -6.500  -0.5117   0.02521   0.01666  -0.0525   1.0000   0.1228
  -6.250  -0.4818   0.02398   0.01513  -0.0531   1.0000   0.1314
  -6.000  -0.4532   0.02267   0.01394  -0.0535   1.0000   0.1399
  -5.750  -0.4237   0.02152   0.01273  -0.0540   1.0000   0.1511
  -5.500  -0.3940   0.02057   0.01178  -0.0544   1.0000   0.1650
  -5.250  -0.3646   0.01974   0.01101  -0.0549   1.0000   0.1822
  -5.000  -0.3350   0.01899   0.01034  -0.0553   1.0000   0.2029
  -4.750  -0.3050   0.01837   0.00971  -0.0559   1.0000   0.2263
  -4.500  -0.2759   0.01789   0.00940  -0.0562   1.0000   0.2480
  -4.250  -0.2463   0.01750   0.00907  -0.0566   1.0000   0.2706
  -4.000  -0.2169   0.01714   0.00876  -0.0571   1.0000   0.2930
  -3.750  -0.1883   0.01684   0.00862  -0.0574   1.0000   0.3143
  -3.500  -0.1603   0.01661   0.00849  -0.0577   1.0000   0.3370
  -3.250  -0.1344   0.01643   0.00836  -0.0578   1.0000   0.3601
  -3.000  -0.1161   0.01640   0.00849  -0.0570   1.0000   0.3789
  -2.750  -0.0933   0.01648   0.00870  -0.0573   0.9968   0.3990
  -2.500  -0.0374   0.01612   0.00839  -0.0629   0.9829   0.4312
  -2.250   0.0136   0.01581   0.00819  -0.0673   0.9640   0.4613
  -2.000   0.0587   0.01558   0.00804  -0.0703   0.9421   0.4890
  -1.750   0.0928   0.01553   0.00800  -0.0709   0.9153   0.5156
  -1.500   0.1190   0.01554   0.00805  -0.0696   0.8901   0.5384
  -1.250   0.1425   0.01555   0.00810  -0.0678   0.8641   0.5622
  -1.000   0.1653   0.01556   0.00811  -0.0658   0.8413   0.5882
  -0.750   0.1884   0.01553   0.00812  -0.0639   0.8191   0.6150
  -0.500   0.2124   0.01547   0.00811  -0.0624   0.7980   0.6431
  -0.250   0.2371   0.01541   0.00809  -0.0610   0.7787   0.6734
   0.000   0.2614   0.01533   0.00807  -0.0595   0.7605   0.7070
   0.250   0.2847   0.01522   0.00805  -0.0577   0.7431   0.7460
   0.500   0.3058   0.01504   0.00801  -0.0554   0.7268   0.7984
   0.750   0.3209   0.01462   0.00779  -0.0514   0.7121   0.8892
   1.000   0.3558   0.01454   0.00756  -0.0528   0.6949   1.0000
   1.250   0.3876   0.01481   0.00760  -0.0536   0.6788   1.0000
   1.500   0.4178   0.01508   0.00769  -0.0538   0.6633   1.0000
   1.750   0.4474   0.01535   0.00780  -0.0540   0.6477   1.0000
   2.000   0.4770   0.01562   0.00796  -0.0541   0.6315   1.0000
   2.250   0.5063   0.01590   0.00814  -0.0541   0.6155   1.0000
   2.500   0.5354   0.01618   0.00833  -0.0541   0.5994   1.0000
   2.750   0.5644   0.01647   0.00853  -0.0541   0.5835   1.0000
   3.000   0.5932   0.01676   0.00875  -0.0540   0.5679   1.0000
   3.250   0.6220   0.01706   0.00897  -0.0539   0.5527   1.0000
   3.500   0.6507   0.01737   0.00921  -0.0537   0.5378   1.0000
   3.750   0.6792   0.01767   0.00944  -0.0536   0.5231   1.0000
   4.000   0.7076   0.01798   0.00965  -0.0533   0.5088   1.0000
   4.250   0.7362   0.01829   0.00994  -0.0532   0.4931   1.0000
   4.500   0.7646   0.01862   0.01027  -0.0531   0.4773   1.0000
   4.750   0.7929   0.01895   0.01059  -0.0530   0.4615   1.0000
   5.000   0.8210   0.01928   0.01090  -0.0528   0.4454   1.0000
   5.250   0.8490   0.01962   0.01123  -0.0526   0.4289   1.0000
   5.500   0.8767   0.01994   0.01153  -0.0524   0.4119   1.0000
   5.750   0.9042   0.02026   0.01182  -0.0521   0.3943   1.0000
   6.000   0.9315   0.02058   0.01207  -0.0518   0.3760   1.0000
   6.250   0.9585   0.02093   0.01237  -0.0515   0.3567   1.0000
   6.500   0.9851   0.02129   0.01280  -0.0512   0.3355   1.0000
   6.750   1.0115   0.02167   0.01315  -0.0509   0.3146   1.0000
   7.000   1.0374   0.02210   0.01349  -0.0505   0.2944   1.0000
   7.250   1.0629   0.02264   0.01404  -0.0502   0.2728   1.0000
   7.500   1.0878   0.02326   0.01465  -0.0499   0.2507   1.0000
   8.000   1.1355   0.02476   0.01608  -0.0491   0.2058   1.0000
   8.250   1.1576   0.02584   0.01697  -0.0486   0.1847   1.0000
   8.500   1.1790   0.02714   0.01828  -0.0479   0.1637   1.0000
   8.750   1.1999   0.02854   0.01962  -0.0473   0.1466   1.0000
   9.000   1.2205   0.03004   0.02110  -0.0465   0.1331   1.0000
   9.250   1.2410   0.03162   0.02265  -0.0458   0.1226   1.0000
   9.500   1.2619   0.03317   0.02405  -0.0451   0.1140   1.0000
   9.750   1.2815   0.03489   0.02603  -0.0442   0.1067   1.0000
  10.000   1.3032   0.03675   0.02765  -0.0437   0.1007   1.0000
  10.250   1.3203   0.03861   0.02992  -0.0426   0.0958   1.0000
  10.500   1.3396   0.04039   0.03175  -0.0418   0.0916   1.0000
  10.750   1.3586   0.04284   0.03420  -0.0412   0.0880   1.0000
  11.000   1.3702   0.04518   0.03696  -0.0398   0.0848   1.0000
  11.250   1.3832   0.04755   0.03956  -0.0387   0.0822   1.0000
  11.500   1.3973   0.04991   0.04202  -0.0377   0.0799   1.0000
  11.750   1.4123   0.05330   0.04542  -0.0372   0.0778   1.0000
  12.000   1.4065   0.05655   0.04918  -0.0351   0.0768   1.0000
  12.250   1.3954   0.06014   0.05319  -0.0330   0.0759   1.0000
  12.500   1.3768   0.06380   0.05719  -0.0306   0.0754   1.0000
  12.750   1.3537   0.06819   0.06188  -0.0296   0.0751   1.0000
  13.000   1.3269   0.07371   0.06770  -0.0305   0.0752   1.0000
  13.250   1.2952   0.08063   0.07491  -0.0332   0.0755   1.0000
  13.500   1.2587   0.08922   0.08375  -0.0378   0.0761   1.0000
  13.750   1.2187   0.09952   0.09427  -0.0441   0.0769   1.0000
  14.000   1.1802   0.11088   0.10578  -0.0509   0.0778   1.0000
<< Back to OAF117 AIRFOIL (oaf117-il)

Polar data table (+)

Polar graphs


<< Back to OAF117 AIRFOIL (oaf117-il)