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

E212 (10.55%) (e212-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: E212 (10.55%) (e212-il)
Reynolds number: 100,000
Max Cl/Cd: 58.75 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e212-il-100000-n5.txt
Download as CSV file: xf-e212-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E212  (10.55%)                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4625   0.09346   0.08826  -0.0406   1.0000   0.0247
  -9.750  -0.4976   0.08040   0.07529  -0.0466   1.0000   0.0236
  -9.500  -0.5136   0.07403   0.06897  -0.0493   1.0000   0.0233
  -9.250  -0.5365   0.06706   0.06203  -0.0529   1.0000   0.0231
  -9.000  -0.5618   0.06161   0.05663  -0.0553   1.0000   0.0228
  -8.750  -0.5866   0.05652   0.05148  -0.0584   1.0000   0.0226
  -8.500  -0.6032   0.05149   0.04623  -0.0607   1.0000   0.0225
  -8.250  -0.6107   0.04686   0.04128  -0.0621   1.0000   0.0225
  -8.000  -0.6094   0.04264   0.03666  -0.0631   1.0000   0.0225
  -7.750  -0.6008   0.03885   0.03239  -0.0638   1.0000   0.0226
  -7.500  -0.5868   0.03553   0.02855  -0.0644   1.0000   0.0232
  -7.250  -0.5690   0.03288   0.02544  -0.0647   1.0000   0.0239
  -7.000  -0.5502   0.03111   0.02357  -0.0647   1.0000   0.0249
  -6.750  -0.5294   0.02944   0.02168  -0.0647   1.0000   0.0258
  -6.500  -0.5074   0.02772   0.01972  -0.0646   1.0000   0.0265
  -6.250  -0.4844   0.02617   0.01792  -0.0645   1.0000   0.0274
  -6.000  -0.4608   0.02484   0.01641  -0.0643   1.0000   0.0283
  -5.750  -0.4367   0.02368   0.01506  -0.0642   1.0000   0.0294
  -5.500  -0.4080   0.02251   0.01381  -0.0652   0.9988   0.0312
  -5.250  -0.3722   0.02164   0.01282  -0.0675   0.9955   0.0350
  -5.000  -0.3366   0.02069   0.01177  -0.0698   0.9920   0.0412
  -4.750  -0.3001   0.01969   0.01072  -0.0721   0.9884   0.0538
  -4.500  -0.2627   0.01875   0.00992  -0.0748   0.9853   0.0919
  -4.250  -0.2284   0.01806   0.00951  -0.0770   0.9809   0.1550
  -4.000  -0.1927   0.01760   0.00926  -0.0793   0.9763   0.2149
  -3.750  -0.1562   0.01731   0.00901  -0.0815   0.9721   0.2576
  -3.500  -0.1232   0.01703   0.00877  -0.0828   0.9661   0.2964
  -3.250  -0.0860   0.01683   0.00860  -0.0849   0.9616   0.3363
  -3.000  -0.0544   0.01662   0.00844  -0.0859   0.9547   0.3710
  -2.750  -0.0181   0.01646   0.00828  -0.0877   0.9496   0.4067
  -2.500   0.0136   0.01630   0.00815  -0.0885   0.9425   0.4397
  -2.250   0.0488   0.01614   0.00803  -0.0900   0.9366   0.4729
  -2.000   0.0812   0.01600   0.00791  -0.0908   0.9295   0.5038
  -1.750   0.1155   0.01585   0.00780  -0.0920   0.9229   0.5344
  -1.500   0.1484   0.01570   0.00770  -0.0928   0.9159   0.5635
  -1.250   0.1816   0.01555   0.00759  -0.0936   0.9086   0.5921
  -1.000   0.2146   0.01539   0.00748  -0.0944   0.9015   0.6203
  -0.750   0.2476   0.01522   0.00735  -0.0950   0.8939   0.6483
  -0.500   0.2798   0.01506   0.00724  -0.0955   0.8858   0.6757
  -0.250   0.3134   0.01488   0.00712  -0.0962   0.8782   0.7029
   0.000   0.3433   0.01474   0.00704  -0.0961   0.8684   0.7294
   0.250   0.3788   0.01452   0.00687  -0.0970   0.8612   0.7571
   0.500   0.4059   0.01440   0.00680  -0.0964   0.8494   0.7839
   0.750   0.4338   0.01423   0.00670  -0.0957   0.8380   0.8111
   1.000   0.4628   0.01403   0.00655  -0.0953   0.8269   0.8398
   1.250   0.4921   0.01380   0.00636  -0.0948   0.8154   0.8710
   1.500   0.5189   0.01359   0.00621  -0.0939   0.8013   0.9085
   1.750   0.5548   0.01336   0.00601  -0.0951   0.7865   0.9680
   2.000   0.5913   0.01330   0.00590  -0.0967   0.7704   1.0000
   2.250   0.6241   0.01335   0.00591  -0.0976   0.7525   1.0000
   2.500   0.6565   0.01341   0.00594  -0.0983   0.7336   1.0000
   2.750   0.6885   0.01350   0.00597  -0.0989   0.7140   1.0000
   3.000   0.7198   0.01362   0.00605  -0.0994   0.6936   1.0000
   3.250   0.7486   0.01379   0.00618  -0.0994   0.6709   1.0000
   3.500   0.7777   0.01398   0.00630  -0.0995   0.6479   1.0000
   3.750   0.8049   0.01421   0.00650  -0.0992   0.6230   1.0000
   4.000   0.8316   0.01447   0.00670  -0.0988   0.5976   1.0000
   4.250   0.8578   0.01477   0.00693  -0.0983   0.5716   1.0000
   4.500   0.8829   0.01510   0.00719  -0.0977   0.5441   1.0000
   4.750   0.9074   0.01545   0.00752  -0.0969   0.5159   1.0000
   5.000   0.9312   0.01585   0.00785  -0.0961   0.4866   1.0000
   5.250   0.9543   0.01629   0.00822  -0.0951   0.4566   1.0000
   5.500   0.9768   0.01676   0.00864  -0.0941   0.4262   1.0000
   5.750   0.9987   0.01728   0.00909  -0.0930   0.3954   1.0000
   6.000   1.0201   0.01783   0.00958  -0.0919   0.3644   1.0000
   6.250   1.0407   0.01844   0.01011  -0.0907   0.3328   1.0000
   6.500   1.0607   0.01911   0.01071  -0.0894   0.3009   1.0000
   6.750   1.0802   0.01983   0.01135  -0.0881   0.2688   1.0000
   7.000   1.0992   0.02060   0.01205  -0.0868   0.2378   1.0000
   7.250   1.1176   0.02142   0.01280  -0.0854   0.2083   1.0000
   7.500   1.1352   0.02233   0.01362  -0.0840   0.1787   1.0000
   7.750   1.1515   0.02335   0.01452  -0.0825   0.1466   1.0000
   8.000   1.1663   0.02451   0.01547  -0.0809   0.1141   1.0000
   8.250   1.1811   0.02570   0.01649  -0.0793   0.0864   1.0000
   8.500   1.1961   0.02685   0.01762  -0.0776   0.0668   1.0000
   8.750   1.2082   0.02823   0.01894  -0.0756   0.0516   1.0000
   9.000   1.2175   0.02972   0.02046  -0.0730   0.0409   1.0000
   9.250   1.2255   0.03126   0.02205  -0.0704   0.0345   1.0000
   9.500   1.2330   0.03281   0.02372  -0.0678   0.0305   1.0000
   9.750   1.2403   0.03434   0.02535  -0.0654   0.0275   1.0000
  10.000   1.2448   0.03612   0.02722  -0.0628   0.0256   1.0000
  10.250   1.2503   0.03787   0.02917  -0.0605   0.0243   1.0000
  10.500   1.2550   0.03976   0.03122  -0.0582   0.0233   1.0000
  10.750   1.2590   0.04177   0.03338  -0.0562   0.0224   1.0000
  11.000   1.2624   0.04391   0.03567  -0.0544   0.0217   1.0000
  11.250   1.2650   0.04622   0.03810  -0.0527   0.0212   1.0000
  11.500   1.2666   0.04878   0.04077  -0.0513   0.0207   1.0000
  11.750   1.2701   0.05141   0.04356  -0.0500   0.0203   1.0000
  12.000   1.2736   0.05417   0.04655  -0.0489   0.0199   1.0000
  12.250   1.2755   0.05715   0.04979  -0.0479   0.0195   1.0000
  12.500   1.2752   0.06039   0.05327  -0.0473   0.0190   1.0000
  12.750   1.2728   0.06391   0.05703  -0.0470   0.0186   1.0000
  13.000   1.2685   0.06773   0.06109  -0.0470   0.0182   1.0000
  13.250   1.2621   0.07190   0.06549  -0.0476   0.0178   1.0000
  13.500   1.2542   0.07646   0.07028  -0.0486   0.0176   1.0000
  13.750   1.2446   0.08146   0.07549  -0.0501   0.0173   1.0000
  14.000   1.2332   0.08696   0.08122  -0.0523   0.0172   1.0000
  14.250   1.2201   0.09311   0.08762  -0.0551   0.0171   1.0000
  14.500   1.2048   0.10005   0.09480  -0.0588   0.0171   1.0000
  14.750   1.1866   0.10806   0.10304  -0.0635   0.0171   1.0000
  15.000   1.1644   0.11764   0.11288  -0.0695   0.0173   1.0000
  15.250   1.1354   0.12988   0.12538  -0.0777   0.0178   1.0000
  15.500   1.0937   0.14748   0.14324  -0.0895   0.0188   1.0000
<< Back to E212 (10.55%) (e212-il)

Polar data table (+)

Polar graphs


<< Back to E212 (10.55%) (e212-il)