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

FX 60-100 AIRFOIL (fx60100-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: FX 60-100 AIRFOIL (fx60100-il)
Reynolds number: 50,000
Max Cl/Cd: 40.54 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx60100-il-50000.txt
Download as CSV file: xf-fx60100-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 60-100 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
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.3758   0.12290   0.11589  -0.0269   1.0000   0.2065
  -9.500  -0.3494   0.11654   0.10951  -0.0253   1.0000   0.2141
  -9.250  -0.3692   0.11706   0.11018  -0.0270   1.0000   0.2207
  -9.000  -0.3426   0.11095   0.10405  -0.0253   1.0000   0.2285
  -8.750  -0.3605   0.11083   0.10410  -0.0264   1.0000   0.2352
  -8.500  -0.3375   0.10554   0.09878  -0.0248   1.0000   0.2438
  -8.250  -0.3489   0.10430   0.09769  -0.0250   1.0000   0.2505
  -8.000  -0.3349   0.10050   0.09388  -0.0238   1.0000   0.2593
  -7.750  -0.3403   0.09838   0.09189  -0.0232   1.0000   0.2664
  -7.500  -0.3445   0.09675   0.09037  -0.0219   1.0000   0.2767
  -7.250  -0.3349   0.09318   0.08686  -0.0206   1.0000   0.2837
  -7.000  -0.3677   0.09424   0.08815  -0.0181   1.0000   0.2921
  -6.750  -0.3425   0.08922   0.08312  -0.0168   1.0000   0.3023
  -6.500  -0.3637   0.08853   0.08262  -0.0138   1.0000   0.3084
  -6.250  -0.3576   0.08561   0.07975  -0.0116   1.0000   0.3165
  -5.750  -0.4087   0.05862   0.05241  -0.0572   1.0000   0.1231
  -5.500  -0.3880   0.05140   0.04473  -0.0629   1.0000   0.1083
  -5.250  -0.3677   0.04666   0.03981  -0.0651   1.0000   0.1028
  -5.000  -0.3284   0.03999   0.03183  -0.0728   1.0000   0.0953
  -4.750  -0.2983   0.03655   0.02778  -0.0750   1.0000   0.0958
  -4.500  -0.2713   0.03346   0.02445  -0.0760   1.0000   0.0982
  -4.250  -0.2429   0.03105   0.02173  -0.0766   1.0000   0.1006
  -4.000  -0.2142   0.02898   0.01932  -0.0767   1.0000   0.1039
  -3.750  -0.1860   0.02726   0.01724  -0.0763   1.0000   0.1099
  -3.500  -0.1608   0.02572   0.01582  -0.0759   1.0000   0.1258
  -3.250  -0.1318   0.02399   0.01426  -0.0761   1.0000   0.1632
  -3.000  -0.1110   0.02348   0.01591  -0.0729   1.0000   0.4881
  -2.750  -0.1049   0.02460   0.01708  -0.0669   1.0000   0.5521
  -2.500  -0.0969   0.02512   0.01761  -0.0617   1.0000   0.5986
  -2.250  -0.0865   0.02521   0.01764  -0.0575   1.0000   0.6356
  -2.000  -0.0730   0.02511   0.01744  -0.0545   1.0000   0.6718
  -1.750  -0.0586   0.02488   0.01711  -0.0519   1.0000   0.7026
  -1.500  -0.0420   0.02464   0.01672  -0.0501   1.0000   0.7318
  -1.250  -0.0244   0.02440   0.01634  -0.0488   1.0000   0.7600
  -1.000  -0.0063   0.02419   0.01601  -0.0477   1.0000   0.7876
  -0.750   0.0110   0.02398   0.01572  -0.0466   1.0000   0.8159
  -0.500   0.0262   0.02377   0.01546  -0.0451   1.0000   0.8465
  -0.250   0.0409   0.02354   0.01522  -0.0438   1.0000   0.8825
   0.000   0.0596   0.02324   0.01500  -0.0439   1.0000   0.9452
   0.250   0.0919   0.02360   0.01518  -0.0486   1.0000   1.0000
   0.500   0.1315   0.02443   0.01575  -0.0543   1.0000   1.0000
   0.750   0.1666   0.02534   0.01642  -0.0589   0.9999   1.0000
   1.000   0.2322   0.02693   0.01773  -0.0686   0.9838   1.0000
   1.250   0.2864   0.02822   0.01880  -0.0757   0.9673   1.0000
   1.500   0.3351   0.02939   0.01982  -0.0813   0.9510   1.0000
   1.750   0.3809   0.03049   0.02081  -0.0862   0.9346   1.0000
   2.000   0.4275   0.03151   0.02174  -0.0907   0.9172   1.0000
   2.250   0.4756   0.03242   0.02261  -0.0951   0.8999   1.0000
   2.500   0.5165   0.03324   0.02343  -0.0982   0.8823   1.0000
   2.750   0.5517   0.03406   0.02426  -0.1003   0.8643   1.0000
   3.000   0.5902   0.03478   0.02502  -0.1026   0.8468   1.0000
   3.250   0.6348   0.03525   0.02555  -0.1053   0.8289   1.0000
   3.500   0.6781   0.03556   0.02599  -0.1074   0.8106   1.0000
   3.750   0.7094   0.03609   0.02662  -0.1080   0.7905   1.0000
   4.000   0.7528   0.03614   0.02680  -0.1097   0.7725   1.0000
   4.250   0.8099   0.03521   0.02613  -0.1119   0.7549   1.0000
   4.500   0.8409   0.03513   0.02618  -0.1112   0.7320   1.0000
   4.750   0.8949   0.03342   0.02474  -0.1116   0.7127   1.0000
   5.000   0.9305   0.03260   0.02409  -0.1103   0.6884   1.0000
   5.250   0.9744   0.03095   0.02262  -0.1090   0.6636   1.0000
   5.500   1.0134   0.02956   0.02139  -0.1073   0.6354   1.0000
   5.750   1.0465   0.02866   0.02057  -0.1052   0.6032   1.0000
   6.000   1.0803   0.02782   0.01971  -0.1031   0.5675   1.0000
   6.250   1.1057   0.02753   0.01935  -0.1004   0.5237   1.0000
   6.500   1.1247   0.02774   0.01943  -0.0973   0.4737   1.0000
   6.750   1.1445   0.02823   0.01961  -0.0945   0.4217   1.0000
   7.000   1.1614   0.02942   0.02058  -0.0919   0.3699   1.0000
   7.250   1.1795   0.03088   0.02169  -0.0897   0.3233   1.0000
   7.500   1.1967   0.03256   0.02316  -0.0877   0.2835   1.0000
   7.750   1.2169   0.03444   0.02485  -0.0862   0.2514   1.0000
   8.000   1.2364   0.03646   0.02683  -0.0849   0.2256   1.0000
   8.250   1.2610   0.03883   0.02906  -0.0842   0.2063   1.0000
   8.500   1.2796   0.04113   0.03157  -0.0829   0.1901   1.0000
   8.750   1.3014   0.04373   0.03426  -0.0820   0.1775   1.0000
   9.000   1.3193   0.04678   0.03772  -0.0807   0.1690   1.0000
   9.250   1.3373   0.05037   0.04159  -0.0796   0.1630   1.0000
   9.500   1.3440   0.05411   0.04599  -0.0775   0.1584   1.0000
   9.750   1.3656   0.05727   0.04910  -0.0770   0.1510   1.0000
  10.000   1.3582   0.06140   0.05394  -0.0740   0.1481   1.0000
  10.250   1.3504   0.06592   0.05896  -0.0715   0.1461   1.0000
  10.500   1.3371   0.07085   0.06432  -0.0691   0.1456   1.0000
  10.750   1.3152   0.07606   0.06988  -0.0668   0.1460   1.0000
  11.000   1.2842   0.08127   0.07533  -0.0645   0.1471   1.0000
  11.250   1.2506   0.08722   0.08144  -0.0638   0.1485   1.0000
  11.500   1.2192   0.09422   0.08855  -0.0650   0.1500   1.0000
  11.750   1.1956   0.10187   0.09622  -0.0673   0.1512   1.0000
<< Back to FX 60-100 AIRFOIL (fx60100-il)

Polar data table (+)

Polar graphs


<< Back to FX 60-100 AIRFOIL (fx60100-il)