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EPPLER 855 AIRFOIL (e855-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 855 AIRFOIL (e855-il)
Reynolds number: 50,000
Max Cl/Cd: 5.67 at α=8.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e855-il-50000.txt
Download as CSV file: xf-e855-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 855 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.3183   0.11854   0.11219  -0.0329   1.0000   0.2944
  -9.250  -0.2916   0.11292   0.10656  -0.0317   1.0000   0.3018
  -9.000  -0.3203   0.11296   0.10678  -0.0304   1.0000   0.3094
  -8.750  -0.3001   0.10794   0.10178  -0.0292   1.0000   0.3146
  -8.500  -0.3050   0.10545   0.09939  -0.0278   1.0000   0.3179
  -8.250  -0.5001   0.08451   0.07909  -0.0477   1.0000   0.1304
  -8.000  -0.5165   0.08129   0.07595  -0.0464   1.0000   0.1264
  -7.750  -0.5865   0.07404   0.06849  -0.0502   1.0000   0.1138
  -7.500  -0.5958   0.07062   0.06506  -0.0491   1.0000   0.1129
  -7.250  -0.6066   0.06677   0.06109  -0.0487   1.0000   0.1119
  -7.000  -0.6130   0.06283   0.05696  -0.0487   1.0000   0.1105
  -6.750  -0.6156   0.05851   0.05231  -0.0493   1.0000   0.1097
  -6.500  -0.6119   0.05428   0.04763  -0.0500   1.0000   0.1093
  -6.250  -0.6021   0.05040   0.04325  -0.0506   1.0000   0.1095
  -6.000  -0.5883   0.04707   0.03930  -0.0512   1.0000   0.1122
  -5.750  -0.5726   0.04417   0.03600  -0.0513   1.0000   0.1170
  -5.500  -0.5560   0.04218   0.03387  -0.0508   1.0000   0.1229
  -5.250  -0.5373   0.04003   0.03130  -0.0507   1.0000   0.1317
  -5.000  -0.4957   0.03811   0.02924  -0.0540   0.9920   0.1492
  -4.750  -0.4531   0.03647   0.02749  -0.0572   0.9834   0.1766
  -4.500  -0.4172   0.03495   0.02619  -0.0593   0.9743   0.2152
  -4.250  -0.3834   0.03350   0.02537  -0.0612   0.9654   0.2838
  -4.000  -0.3556   0.03422   0.02718  -0.0601   0.9558   0.4252
  -3.750  -0.3337   0.03671   0.02981  -0.0561   0.9458   0.5130
  -3.500  -0.3161   0.03838   0.03147  -0.0519   0.9361   0.5609
  -3.250  -0.3017   0.03947   0.03249  -0.0479   0.9268   0.5963
  -3.000  -0.2785   0.04039   0.03323  -0.0457   0.9186   0.6332
  -2.750  -0.2682   0.04088   0.03368  -0.0415   0.9099   0.6577
  -2.500  -0.2501   0.04132   0.03396  -0.0390   0.9023   0.6854
  -2.250  -0.2349   0.04153   0.03402  -0.0368   0.8945   0.7118
  -2.000  -0.2206   0.04178   0.03418  -0.0336   0.8874   0.7341
  -1.750  -0.2065   0.04184   0.03410  -0.0317   0.8800   0.7558
  -1.500  -0.1844   0.04195   0.03405  -0.0310   0.8737   0.7759
  -1.250  -0.1737   0.04194   0.03392  -0.0291   0.8667   0.7917
  -1.000  -0.1440   0.04209   0.03390  -0.0301   0.8605   0.8087
  -0.750  -0.1345   0.04217   0.03387  -0.0288   0.8545   0.8221
  -0.500  -0.1124   0.04235   0.03390  -0.0292   0.8486   0.8367
  -0.250  -0.0879   0.04263   0.03404  -0.0300   0.8428   0.8512
   0.000  -0.0759   0.04289   0.03422  -0.0292   0.8378   0.8654
   0.250  -0.0515   0.04326   0.03450  -0.0301   0.8326   0.8812
   0.500  -0.0268   0.04373   0.03488  -0.0311   0.8276   0.8986
   0.750  -0.0085   0.04426   0.03538  -0.0318   0.8231   0.9172
   1.000   0.0322   0.04512   0.03618  -0.0364   0.8178   0.9402
   1.250   0.0866   0.04635   0.03731  -0.0440   0.8114   0.9665
   1.500   0.1131   0.04727   0.03817  -0.0480   0.8075   1.0000
   1.750   0.1302   0.04824   0.03901  -0.0503   0.8055   1.0000
   2.000   0.1605   0.04949   0.04012  -0.0544   0.8021   1.0000
   2.250   0.1952   0.05095   0.04143  -0.0589   0.7978   1.0000
   2.500   0.2160   0.05254   0.04289  -0.0616   0.7976   1.0000
   2.750   0.2383   0.05428   0.04450  -0.0643   0.7982   1.0000
   3.000   0.1006   0.05553   0.04599  -0.0488   0.9456   1.0000
   3.250   0.1301   0.05727   0.04758  -0.0521   0.9366   1.0000
   3.500   0.1539   0.05892   0.04910  -0.0542   0.9297   1.0000
   3.750   0.1869   0.06128   0.05131  -0.0577   0.9211   1.0000
   4.000   0.2045   0.06249   0.05244  -0.0586   0.9117   1.0000
   4.250   0.2387   0.06539   0.05522  -0.0621   0.9046   1.0000
   4.500   0.2569   0.06656   0.05634  -0.0629   0.8930   1.0000
   4.750   0.2750   0.06815   0.05787  -0.0636   0.8833   1.0000
   5.000   0.3117   0.07144   0.06108  -0.0673   0.8751   1.0000
   5.250   0.3250   0.07239   0.06202  -0.0672   0.8627   1.0000
   5.500   0.3409   0.07398   0.06359  -0.0676   0.8518   1.0000
   5.750   0.3681   0.07667   0.06625  -0.0697   0.8431   1.0000
   6.000   0.3921   0.07879   0.06838  -0.0712   0.8309   1.0000
   6.250   0.4050   0.08023   0.06983  -0.0711   0.8185   1.0000
   6.500   0.4213   0.08220   0.07181  -0.0716   0.8074   1.0000
   6.750   0.4509   0.08531   0.07493  -0.0739   0.7980   1.0000
   7.000   0.4716   0.08742   0.07709  -0.0749   0.7847   1.0000
   7.250   0.4827   0.08900   0.07871  -0.0746   0.7715   1.0000
   7.500   0.4960   0.09098   0.08073  -0.0748   0.7589   1.0000
   7.750   0.5125   0.09335   0.08315  -0.0754   0.7475   1.0000
   8.000   0.5393   0.09652   0.08639  -0.0772   0.7364   1.0000
   8.250   0.5628   0.09926   0.08921  -0.0784   0.7228   1.0000
   8.500   0.5729   0.10110   0.09112  -0.0783   0.7085   1.0000
   8.750   0.5834   0.10318   0.09329  -0.0782   0.6946   1.0000
   9.000   0.5956   0.10552   0.09571  -0.0784   0.6806   1.0000
   9.250   0.6084   0.10799   0.09826  -0.0787   0.6668   1.0000
   9.500   0.6219   0.11060   0.10096  -0.0791   0.6527   1.0000
   9.750   0.6357   0.11329   0.10375  -0.0796   0.6386   1.0000
  10.000   0.6496   0.11604   0.10662  -0.0800   0.6242   1.0000
  10.250   0.6632   0.11878   0.10945  -0.0805   0.6089   1.0000
  10.500   0.6762   0.12158   0.11236  -0.0809   0.5937   1.0000
  10.750   0.6891   0.12427   0.11516  -0.0812   0.5770   1.0000
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