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

WORTMANN FX M2 AIRFOIL (fxm2-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: WORTMANN FX M2 AIRFOIL (fxm2-il)
Reynolds number: 500,000
Max Cl/Cd: 86.22 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-fxm2-il-500000-n5.txt
Download as CSV file: xf-fxm2-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: WORTMANN FX M2 AIRFOIL                          
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.3150   0.11805   0.11568  -0.0111   1.0000   0.0108
  -9.750  -0.3098   0.11521   0.11286  -0.0116   1.0000   0.0113
  -9.500  -0.3049   0.11235   0.11002  -0.0122   1.0000   0.0115
  -9.250  -0.2927   0.10850   0.10618  -0.0153   0.9985   0.0116
  -9.000  -0.2818   0.10480   0.10248  -0.0180   0.9936   0.0117
  -8.750  -0.2725   0.10166   0.09936  -0.0197   0.9885   0.0117
  -8.500  -0.2622   0.09798   0.09569  -0.0224   0.9793   0.0117
  -8.000  -0.1507   0.08415   0.08068  -0.0534   0.6833   0.0117
  -7.750  -0.1453   0.08259   0.07888  -0.0524   0.6186   0.0106
  -7.500  -0.1406   0.08016   0.07634  -0.0529   0.5914   0.0110
  -7.250  -0.1376   0.07744   0.07358  -0.0538   0.5648   0.0122
  -7.000  -0.1348   0.07515   0.07121  -0.0540   0.5444   0.0122
  -6.750  -0.1265   0.07382   0.06985  -0.0540   0.5302   0.0131
  -6.500  -0.1179   0.07078   0.06679  -0.0579   0.5168   0.0148
  -6.250  -0.1116   0.06627   0.06222  -0.0609   0.4945   0.0117
  -6.000  -0.0990   0.06360   0.05947  -0.0631   0.4796   0.0115
  -5.750  -0.0843   0.06050   0.05629  -0.0663   0.4638   0.0113
  -5.500  -0.0667   0.05714   0.05288  -0.0700   0.4466   0.0111
  -5.250  -0.0477   0.05353   0.04907  -0.0740   0.4155   0.0111
  -5.000  -0.0247   0.04942   0.04485  -0.0785   0.4057   0.0112
  -4.750   0.0066   0.04188   0.03707  -0.0863   0.4003   0.0119
  -4.500   0.0314   0.03966   0.03472  -0.0882   0.3765   0.0121
  -4.250   0.0561   0.03788   0.03274  -0.0895   0.3531   0.0125
  -4.000   0.0838   0.03528   0.02996  -0.0914   0.3466   0.0129
  -3.750   0.1134   0.03237   0.02684  -0.0933   0.3409   0.0139
  -3.500   0.1492   0.02655   0.02054  -0.0962   0.3362   0.0152
  -3.250   0.1765   0.02538   0.01923  -0.0966   0.3270   0.0155
  -3.000   0.2044   0.02401   0.01764  -0.0971   0.3012   0.0161
  -2.500   0.2675   0.01821   0.01089  -0.0983   0.2879   0.0187
  -2.250   0.2953   0.01760   0.01012  -0.0983   0.2818   0.0190
  -2.000   0.3234   0.01696   0.00934  -0.0983   0.2760   0.0196
  -1.750   0.3523   0.01603   0.00821  -0.0983   0.2688   0.0204
  -1.500   0.3817   0.01490   0.00682  -0.0982   0.2549   0.0215
  -1.250   0.4097   0.01426   0.00591  -0.0980   0.2416   0.0228
  -1.000   0.4373   0.01401   0.00555  -0.0978   0.2368   0.0234
  -0.750   0.4651   0.01371   0.00515  -0.0976   0.2324   0.0241
  -0.500   0.4929   0.01339   0.00473  -0.0973   0.2283   0.0250
  -0.250   0.5206   0.01313   0.00435  -0.0970   0.2242   0.0259
   0.000   0.5483   0.01296   0.00409  -0.0968   0.2205   0.0271
   0.250   0.5762   0.01274   0.00390  -0.0966   0.2164   0.0281
   0.500   0.6040   0.01260   0.00374  -0.0964   0.2119   0.0290
   0.750   0.6320   0.01247   0.00358  -0.0962   0.2086   0.0300
   1.000   0.6602   0.01234   0.00344  -0.0961   0.2038   0.0310
   1.250   0.6880   0.01232   0.00338  -0.0958   0.1938   0.0322
   1.500   0.7159   0.01224   0.00325  -0.0958   0.1847   0.0338
   1.750   0.7436   0.01225   0.00321  -0.0956   0.1791   0.0350
   2.000   0.7711   0.01231   0.00320  -0.0953   0.1729   0.0363
   2.250   0.7987   0.01236   0.00323  -0.0951   0.1691   0.0376
   2.500   0.8263   0.01240   0.00325  -0.0949   0.1650   0.0399
   2.750   0.8538   0.01248   0.00331  -0.0947   0.1611   0.0423
   3.000   0.8809   0.01260   0.00340  -0.0944   0.1576   0.0444
   3.250   0.9082   0.01272   0.00350  -0.0941   0.1543   0.0484
   3.500   0.9359   0.01274   0.00363  -0.0940   0.1507   0.1107
   3.750   0.9645   0.01254   0.00385  -0.0944   0.1455   0.3704
   4.000   0.9887   0.01176   0.00417  -0.0939   0.1399   0.8467
   4.500   1.0355   0.01201   0.00443  -0.0916   0.1218   1.0000
   4.750   1.0615   0.01234   0.00465  -0.0913   0.1118   1.0000
   5.000   1.0877   0.01264   0.00488  -0.0909   0.1041   1.0000
   5.250   1.1136   0.01295   0.00514  -0.0905   0.0978   1.0000
   5.500   1.1397   0.01324   0.00539  -0.0901   0.0910   1.0000
   5.750   1.1652   0.01359   0.00568  -0.0897   0.0854   1.0000
   6.000   1.1911   0.01389   0.00595  -0.0893   0.0833   1.0000
   6.250   1.2168   0.01419   0.00625  -0.0889   0.0814   1.0000
   6.500   1.2424   0.01449   0.00653  -0.0885   0.0798   1.0000
   6.750   1.2680   0.01479   0.00682  -0.0881   0.0784   1.0000
   7.000   1.2935   0.01509   0.00712  -0.0877   0.0772   1.0000
   7.250   1.3189   0.01538   0.00744  -0.0873   0.0765   1.0000
   7.500   1.3441   0.01568   0.00775  -0.0869   0.0760   1.0000
   7.750   1.3692   0.01599   0.00808  -0.0864   0.0755   1.0000
   8.000   1.3941   0.01630   0.00842  -0.0860   0.0749   1.0000
   8.250   1.4188   0.01663   0.00879  -0.0855   0.0743   1.0000
   8.500   1.4432   0.01698   0.00917  -0.0850   0.0738   1.0000
   8.750   1.4673   0.01734   0.00956  -0.0844   0.0732   1.0000
   9.000   1.4913   0.01771   0.00997  -0.0839   0.0726   1.0000
   9.250   1.5149   0.01810   0.01040  -0.0833   0.0721   1.0000
   9.500   1.5383   0.01849   0.01084  -0.0827   0.0717   1.0000
   9.750   1.5614   0.01890   0.01129  -0.0820   0.0713   1.0000
  10.000   1.5840   0.01934   0.01179  -0.0813   0.0707   1.0000
  10.250   1.6059   0.01984   0.01233  -0.0806   0.0697   1.0000
  10.500   1.6272   0.02037   0.01291  -0.0797   0.0686   1.0000
  10.750   1.6481   0.02090   0.01349  -0.0788   0.0676   1.0000
  11.000   1.6693   0.02138   0.01404  -0.0780   0.0671   1.0000
  11.250   1.6907   0.02180   0.01454  -0.0772   0.0666   1.0000
  11.500   1.7114   0.02227   0.01509  -0.0763   0.0662   1.0000
  11.750   1.7314   0.02276   0.01566  -0.0753   0.0658   1.0000
  12.000   1.7507   0.02328   0.01627  -0.0742   0.0653   1.0000
  12.250   1.7692   0.02383   0.01691  -0.0730   0.0647   1.0000
  12.500   1.7868   0.02440   0.01757  -0.0717   0.0640   1.0000
  12.750   1.8027   0.02500   0.01827  -0.0702   0.0631   1.0000
  13.000   1.8162   0.02564   0.01900  -0.0683   0.0621   1.0000
  13.250   1.8283   0.02633   0.01978  -0.0662   0.0610   1.0000
  13.500   1.8393   0.02710   0.02065  -0.0642   0.0596   1.0000
  13.750   1.8494   0.02794   0.02158  -0.0621   0.0585   1.0000
  14.000   1.8590   0.02884   0.02260  -0.0601   0.0572   1.0000
  14.250   1.8676   0.02985   0.02372  -0.0582   0.0551   1.0000
  14.500   1.8745   0.03104   0.02499  -0.0563   0.0516   1.0000
  14.750   1.8786   0.03251   0.02651  -0.0545   0.0505   1.0000
  15.000   1.8798   0.03433   0.02836  -0.0528   0.0488   1.0000
  15.250   1.8780   0.03654   0.03063  -0.0513   0.0469   1.0000
  15.500   1.8749   0.03909   0.03328  -0.0503   0.0459   1.0000
  15.750   1.8695   0.04213   0.03644  -0.0498   0.0453   1.0000
  16.000   1.8610   0.04587   0.04032  -0.0501   0.0447   1.0000
  16.250   1.8493   0.05044   0.04504  -0.0511   0.0442   1.0000
  16.500   1.8341   0.05595   0.05071  -0.0530   0.0437   1.0000
  16.750   1.8150   0.06245   0.05739  -0.0558   0.0434   1.0000
  17.000   1.7921   0.06993   0.06505  -0.0592   0.0433   1.0000
  17.250   1.7658   0.07823   0.07355  -0.0632   0.0432   1.0000
  17.500   1.7369   0.08718   0.08268  -0.0675   0.0433   1.0000
  17.750   1.7064   0.09667   0.09236  -0.0723   0.0433   1.0000
<< Back to WORTMANN FX M2 AIRFOIL (fxm2-il)

Polar data table (+)

Polar graphs


<< Back to WORTMANN FX M2 AIRFOIL (fxm2-il)