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

EPPLER 908 AIRFOIL (e908-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 908 AIRFOIL (e908-il)
Reynolds number: 100,000
Max Cl/Cd: 32.2 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e908-il-100000.txt
Download as CSV file: xf-e908-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 908 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4998   0.10549   0.10106  -0.0297   1.0000   0.0876
  -8.000  -0.5211   0.10373   0.09941  -0.0303   1.0000   0.0883
  -7.750  -0.5425   0.10177   0.09754  -0.0305   1.0000   0.0886
  -7.500  -0.5147   0.09616   0.09190  -0.0257   1.0000   0.0942
  -7.250  -0.5261   0.09389   0.08970  -0.0243   1.0000   0.0966
  -7.000  -0.5403   0.09116   0.08705  -0.0249   1.0000   0.0989
  -6.750  -0.5572   0.08703   0.08291  -0.0321   1.0000   0.1012
  -6.500  -0.5590   0.08235   0.07828  -0.0317   1.0000   0.1037
  -6.250  -0.5524   0.07973   0.07570  -0.0280   1.0000   0.1088
  -6.000  -0.5601   0.07471   0.07038  -0.0373   1.0000   0.1157
  -5.750  -0.5505   0.07139   0.06729  -0.0314   1.0000   0.1201
  -5.500  -0.5459   0.06664   0.06234  -0.0357   1.0000   0.1301
  -5.250  -0.5359   0.06355   0.05918  -0.0339   1.0000   0.1376
  -5.000  -0.5253   0.05958   0.05512  -0.0349   1.0000   0.1477
  -4.750  -0.5117   0.05575   0.05116  -0.0361   1.0000   0.1606
  -4.500  -0.4955   0.05215   0.04742  -0.0370   1.0000   0.1748
  -4.250  -0.4767   0.04870   0.04381  -0.0381   1.0000   0.1903
  -4.000  -0.4004   0.03416   0.02639  -0.0457   1.0000   0.0501
  -3.750  -0.3741   0.03007   0.02199  -0.0460   1.0000   0.0458
  -3.500  -0.3457   0.02738   0.01871  -0.0455   1.0000   0.0420
  -3.250  -0.3187   0.02553   0.01644  -0.0446   1.0000   0.0401
  -3.000  -0.2929   0.02400   0.01469  -0.0438   1.0000   0.0396
  -2.750  -0.2677   0.02266   0.01327  -0.0431   1.0000   0.0407
  -2.500  -0.2430   0.02173   0.01228  -0.0427   1.0000   0.0470
  -2.250  -0.2166   0.02067   0.01119  -0.0424   1.0000   0.0650
  -2.000  -0.1926   0.01768   0.01117  -0.0415   1.0000   0.6628
  -1.750  -0.1986   0.01748   0.01164  -0.0319   1.0000   0.8571
  -1.500  -0.1739   0.01699   0.01097  -0.0308   1.0000   1.0000
  -1.250  -0.1478   0.01718   0.01079  -0.0316   1.0000   1.0000
  -1.000  -0.1224   0.01743   0.01072  -0.0323   1.0000   1.0000
  -0.750  -0.0975   0.01771   0.01073  -0.0328   1.0000   1.0000
  -0.500  -0.0733   0.01803   0.01083  -0.0331   1.0000   1.0000
  -0.250  -0.0497   0.01838   0.01098  -0.0333   1.0000   1.0000
   0.000  -0.0266   0.01876   0.01109  -0.0334   1.0000   1.0000
   0.250  -0.0040   0.01917   0.01135  -0.0335   1.0000   1.0000
   0.500   0.0182   0.01961   0.01166  -0.0335   1.0000   1.0000
   0.750   0.0400   0.02008   0.01203  -0.0334   1.0000   1.0000
   1.000   0.0614   0.02057   0.01244  -0.0333   1.0000   1.0000
   1.250   0.0971   0.02149   0.01328  -0.0360   0.9950   1.0000
   1.500   0.1382   0.02240   0.01415  -0.0398   0.9858   1.0000
   1.750   0.1760   0.02320   0.01492  -0.0428   0.9769   1.0000
   2.000   0.2170   0.02420   0.01590  -0.0464   0.9702   1.0000
   2.250   0.2488   0.02467   0.01640  -0.0482   0.9605   1.0000
   2.500   0.2814   0.02525   0.01702  -0.0501   0.9514   1.0000
   2.750   0.3186   0.02597   0.01787  -0.0527   0.9437   1.0000
   3.000   0.3521   0.02652   0.01852  -0.0547   0.9347   1.0000
   3.250   0.3824   0.02702   0.01913  -0.0560   0.9248   1.0000
   3.500   0.4167   0.02758   0.01984  -0.0580   0.9160   1.0000
   3.750   0.4555   0.02814   0.02059  -0.0607   0.9079   1.0000
   4.000   0.4854   0.02854   0.02120  -0.0617   0.8969   1.0000
   4.250   0.5186   0.02890   0.02194  -0.0632   0.8857   1.0000
   4.500   0.5567   0.02908   0.02247  -0.0651   0.8733   1.0000
   4.750   0.7210   0.02239   0.01216  -0.0680   0.0406   1.0000
   5.000   0.7588   0.02558   0.01522  -0.0695   0.0386   1.0000
   5.250   0.8107   0.02884   0.01871  -0.0728   0.0390   1.0000
   5.500   0.8553   0.03323   0.02338  -0.0749   0.0385   1.0000
   5.750   0.8763   0.03411   0.02475  -0.0724   0.0360   1.0000
   6.000   0.9092   0.03935   0.03022  -0.0730   0.0387   1.0000
   6.250   0.9397   0.04171   0.03361  -0.0689   0.0588   1.0000
   9.250   0.8889   0.08395   0.07972  -0.0267   0.1011   1.0000
   9.500   0.8621   0.08786   0.08374  -0.0246   0.1011   1.0000
   9.750   0.8364   0.09247   0.08844  -0.0240   0.1010   1.0000
<< Back to EPPLER 908 AIRFOIL (e908-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 908 AIRFOIL (e908-il)