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=50,000 Ncrit=9


Details Polar file
Airfoil: OAF117 AIRFOIL (oaf117-il)
Reynolds number: 50,000
Max Cl/Cd: 28.29 at α=8.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-oaf117-il-50000.txt
Download as CSV file: xf-oaf117-il-50000.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 =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4969   0.10103   0.09397   0.0239   1.0000   0.3279
  -7.500  -0.4937   0.09837   0.09135   0.0234   1.0000   0.3409
  -7.250  -0.5601   0.07210   0.06504  -0.0172   1.0000   0.1938
  -7.000  -0.5491   0.06546   0.05829  -0.0221   1.0000   0.1913
  -6.750  -0.5380   0.05727   0.04985  -0.0298   1.0000   0.1899
  -6.500  -0.5210   0.04813   0.04014  -0.0390   1.0000   0.1906
  -6.250  -0.4973   0.04401   0.03580  -0.0417   1.0000   0.1977
  -6.000  -0.4694   0.03917   0.03039  -0.0460   1.0000   0.2070
  -5.750  -0.4430   0.03718   0.02838  -0.0464   1.0000   0.2185
  -5.500  -0.4143   0.03459   0.02555  -0.0480   1.0000   0.2324
  -5.250  -0.3858   0.03266   0.02349  -0.0489   1.0000   0.2499
  -5.000  -0.3565   0.03092   0.02158  -0.0498   1.0000   0.2716
  -4.750  -0.3283   0.02963   0.02025  -0.0501   1.0000   0.2945
  -4.500  -0.3017   0.02885   0.01958  -0.0496   1.0000   0.3180
  -4.250  -0.2733   0.02783   0.01849  -0.0499   1.0000   0.3447
  -4.000  -0.2478   0.02727   0.01807  -0.0490   1.0000   0.3688
  -3.750  -0.2199   0.02640   0.01714  -0.0493   1.0000   0.3959
  -3.500  -0.1957   0.02583   0.01672  -0.0484   1.0000   0.4192
  -3.250  -0.1694   0.02506   0.01592  -0.0486   1.0000   0.4456
  -3.000  -0.1475   0.02445   0.01542  -0.0477   1.0000   0.4681
  -2.750  -0.1290   0.02400   0.01509  -0.0465   1.0000   0.4906
  -2.500  -0.1107   0.02362   0.01472  -0.0458   1.0000   0.5149
  -2.250  -0.0944   0.02340   0.01458  -0.0449   1.0000   0.5376
  -2.000  -0.0779   0.02332   0.01458  -0.0441   1.0000   0.5615
  -1.750  -0.0592   0.02330   0.01460  -0.0441   1.0000   0.5883
  -1.500  -0.0419   0.02342   0.01481  -0.0437   1.0000   0.6152
  -1.250  -0.0229   0.02361   0.01509  -0.0437   1.0000   0.6451
  -1.000   0.0324   0.02338   0.01495  -0.0490   0.9820   0.6937
  -0.750   0.0851   0.02306   0.01473  -0.0537   0.9616   0.7420
  -0.500   0.1345   0.02264   0.01446  -0.0570   0.9430   0.7941
  -0.250   0.1752   0.02211   0.01410  -0.0581   0.9247   0.8564
   0.000   0.2228   0.02149   0.01365  -0.0608   0.9018   0.9680
   0.250   0.2905   0.02147   0.01334  -0.0710   0.8773   1.0000
   0.500   0.3371   0.02177   0.01337  -0.0751   0.8561   1.0000
   0.750   0.3710   0.02228   0.01369  -0.0766   0.8339   1.0000
   1.000   0.4025   0.02280   0.01405  -0.0772   0.8137   1.0000
   1.250   0.4319   0.02334   0.01446  -0.0772   0.7945   1.0000
   1.500   0.4601   0.02389   0.01492  -0.0768   0.7760   1.0000
   1.750   0.4874   0.02448   0.01542  -0.0763   0.7579   1.0000
   2.000   0.5141   0.02507   0.01593  -0.0756   0.7400   1.0000
   2.250   0.5405   0.02568   0.01649  -0.0749   0.7223   1.0000
   2.500   0.5666   0.02629   0.01705  -0.0740   0.7046   1.0000
   2.750   0.5925   0.02692   0.01764  -0.0732   0.6871   1.0000
   3.000   0.6183   0.02754   0.01823  -0.0722   0.6695   1.0000
   3.250   0.6440   0.02817   0.01884  -0.0713   0.6519   1.0000
   3.500   0.6696   0.02881   0.01947  -0.0703   0.6344   1.0000
   3.750   0.6952   0.02945   0.02012  -0.0693   0.6170   1.0000
   4.000   0.7207   0.03012   0.02078  -0.0683   0.5995   1.0000
   4.250   0.7462   0.03078   0.02146  -0.0673   0.5821   1.0000
   4.500   0.7716   0.03144   0.02214  -0.0662   0.5645   1.0000
   4.750   0.7970   0.03210   0.02282  -0.0651   0.5469   1.0000
   5.000   0.8224   0.03271   0.02345  -0.0639   0.5291   1.0000
   5.250   0.8478   0.03333   0.02409  -0.0627   0.5112   1.0000
   5.500   0.8732   0.03386   0.02463  -0.0613   0.4931   1.0000
   5.750   0.8988   0.03437   0.02512  -0.0600   0.4750   1.0000
   6.000   0.9220   0.03555   0.02647  -0.0594   0.4536   1.0000
   6.250   0.9462   0.03631   0.02730  -0.0583   0.4332   1.0000
   6.500   0.9712   0.03679   0.02778  -0.0569   0.4133   1.0000
   6.750   0.9971   0.03695   0.02790  -0.0552   0.3935   1.0000
   7.000   1.0222   0.03730   0.02819  -0.0537   0.3728   1.0000
   7.250   1.0444   0.03822   0.02924  -0.0526   0.3496   1.0000
   7.500   1.0702   0.03825   0.02913  -0.0509   0.3278   1.0000
   7.750   1.0928   0.03902   0.02992  -0.0497   0.3048   1.0000
   8.000   1.1155   0.03971   0.03058  -0.0483   0.2816   1.0000
   8.250   1.1400   0.04030   0.03092  -0.0469   0.2601   1.0000
   8.500   1.1572   0.04226   0.03312  -0.0460   0.2388   1.0000
   8.750   1.1765   0.04395   0.03485  -0.0449   0.2204   1.0000
   9.000   1.1947   0.04598   0.03694  -0.0438   0.2049   1.0000
   9.250   1.2100   0.04858   0.03970  -0.0429   0.1921   1.0000
   9.500   1.2275   0.05102   0.04217  -0.0420   0.1815   1.0000
   9.750   1.2387   0.05409   0.04549  -0.0410   0.1724   1.0000
  10.000   1.2424   0.05823   0.04994  -0.0402   0.1663   1.0000
  10.250   1.2268   0.06416   0.05641  -0.0396   0.1633   1.0000
  10.500   1.1979   0.07112   0.06377  -0.0397   0.1626   1.0000
  10.750   1.1453   0.08015   0.07307  -0.0415   0.1654   1.0000
  11.000   1.0956   0.09166   0.08465  -0.0475   0.1682   1.0000
<< Back to OAF117 AIRFOIL (oaf117-il)

Polar data table (+)

Polar graphs


<< Back to OAF117 AIRFOIL (oaf117-il)