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EPPLER 857 AIRFOIL (e857-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 857 AIRFOIL (e857-il)
Reynolds number: 100,000
Max Cl/Cd: 25.86 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e857-il-100000.txt
Download as CSV file: xf-e857-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 857 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
  ------ -------- --------- --------- -------- -------- --------
 -12.250  -0.2823   0.13961   0.13503  -0.0462   1.0000   0.1325
 -12.000  -0.5155   0.08088   0.07600  -0.0760   1.0000   0.0586
 -11.750  -0.5463   0.07647   0.07157  -0.0763   1.0000   0.0583
 -11.500  -0.5714   0.07324   0.06837  -0.0752   1.0000   0.0579
 -11.250  -0.6037   0.07062   0.06576  -0.0734   1.0000   0.0575
 -11.000  -0.6372   0.06865   0.06381  -0.0709   1.0000   0.0571
 -10.750  -0.6591   0.06421   0.05918  -0.0752   0.9936   0.0566
 -10.500  -0.6747   0.05903   0.05356  -0.0814   0.9789   0.0570
 -10.250  -0.6738   0.05413   0.04813  -0.0856   0.9651   0.0576
 -10.000  -0.6670   0.04958   0.04289  -0.0886   0.9519   0.0586
  -9.750  -0.6364   0.04686   0.04022  -0.0910   0.9426   0.0607
  -9.500  -0.6038   0.04385   0.03693  -0.0943   0.9349   0.0632
  -9.250  -0.5815   0.04091   0.03350  -0.0958   0.9233   0.0667
  -9.000  -0.5512   0.03922   0.03188  -0.0972   0.9124   0.0710
  -8.750  -0.5115   0.03675   0.02924  -0.1004   0.9066   0.0773
  -8.500  -0.4864   0.03509   0.02732  -0.1008   0.8942   0.0843
  -8.250  -0.4420   0.03295   0.02517  -0.1043   0.8897   0.0953
  -8.000  -0.4173   0.03156   0.02381  -0.1043   0.8773   0.1054
  -7.750  -0.3730   0.02938   0.02177  -0.1077   0.8732   0.1217
  -7.500  -0.3477   0.02805   0.02054  -0.1077   0.8610   0.1370
  -7.250  -0.3017   0.02616   0.01871  -0.1115   0.8561   0.1637
  -7.000  -0.2751   0.02482   0.01760  -0.1120   0.8443   0.1899
  -6.750  -0.2293   0.02298   0.01611  -0.1161   0.8381   0.2395
  -6.500  -0.2017   0.02225   0.01571  -0.1165   0.8255   0.2993
  -6.250  -0.1477   0.02253   0.01604  -0.1204   0.8182   0.3645
  -6.000  -0.1174   0.02333   0.01681  -0.1197   0.8043   0.3922
  -5.750  -0.0761   0.02415   0.01753  -0.1209   0.7930   0.4167
  -5.500  -0.0349   0.02487   0.01813  -0.1220   0.7819   0.4368
  -5.250  -0.0078   0.02554   0.01875  -0.1208   0.7688   0.4506
  -5.000   0.0315   0.02590   0.01891  -0.1220   0.7588   0.4665
  -4.750   0.0541   0.02612   0.01899  -0.1207   0.7461   0.4799
  -4.500   0.0829   0.02660   0.01944  -0.1197   0.7350   0.4889
  -4.250   0.1117   0.02651   0.01910  -0.1200   0.7248   0.5030
  -4.000   0.1349   0.02721   0.01985  -0.1176   0.7140   0.5117
  -3.750   0.1645   0.02726   0.01971  -0.1177   0.7053   0.5253
  -3.500   0.1841   0.02760   0.02004  -0.1155   0.6952   0.5343
  -3.250   0.2117   0.02727   0.01949  -0.1159   0.6867   0.5445
  -3.000   0.2324   0.02746   0.01965  -0.1141   0.6777   0.5515
  -2.750   0.2566   0.02705   0.01902  -0.1145   0.6693   0.5622
  -2.500   0.2835   0.02719   0.01911  -0.1137   0.6623   0.5681
  -2.250   0.3013   0.02713   0.01899  -0.1123   0.6539   0.5757
  -2.000   0.3321   0.02682   0.01847  -0.1134   0.6477   0.5836
  -1.750   0.3497   0.02694   0.01860  -0.1115   0.6401   0.5888
  -1.500   0.3732   0.02682   0.01837  -0.1112   0.6329   0.5960
  -1.250   0.4070   0.02658   0.01790  -0.1128   0.6276   0.6031
  -1.000   0.4197   0.02678   0.01818  -0.1102   0.6203   0.6078
  -0.750   0.4434   0.02678   0.01811  -0.1099   0.6143   0.6141
  -0.500   0.4760   0.02665   0.01777  -0.1113   0.6095   0.6216
  -0.250   0.4923   0.02689   0.01807  -0.1094   0.6036   0.6265
   0.000   0.5113   0.02704   0.01822  -0.1082   0.5973   0.6322
   0.250   0.5421   0.02700   0.01800  -0.1094   0.5924   0.6400
   0.500   0.5707   0.02711   0.01806  -0.1095   0.5883   0.6457
   0.750   0.5789   0.02760   0.01867  -0.1066   0.5824   0.6515
   1.000   0.6018   0.02784   0.01883  -0.1064   0.5774   0.6588
   1.250   0.6302   0.02790   0.01887  -0.1065   0.5734   0.6648
   1.500   0.6603   0.02811   0.01898  -0.1072   0.5696   0.6727
   1.750   0.6628   0.02885   0.01985  -0.1036   0.5636   0.6794
   2.000   0.6808   0.02922   0.02027  -0.1021   0.5590   0.6857
   2.250   0.7090   0.02943   0.02043  -0.1025   0.5554   0.6943
   2.500   0.7426   0.02957   0.02050  -0.1037   0.5524   0.7028
   2.750   0.7382   0.03074   0.02185  -0.0990   0.5475   0.7103
   3.000   0.7413   0.03171   0.02293  -0.0957   0.5423   0.7187
   3.250   0.7645   0.03201   0.02325  -0.0952   0.5385   0.7279
   3.500   0.7977   0.03209   0.02328  -0.0962   0.5355   0.7392
   3.750   0.8346   0.03228   0.02340  -0.0978   0.5330   0.7523
   4.000   0.7351   0.03683   0.02843  -0.0805   0.5249   0.7566
   4.250   0.7254   0.03854   0.03024  -0.0759   0.5202   0.7680
   4.500   0.7759   0.03781   0.02950  -0.0788   0.5181   0.7855
   4.750   0.8346   0.03702   0.02867  -0.0830   0.5162   0.8069
   6.750   0.3912   0.09614   0.08913  -0.0617   0.5228   1.0000
   7.000   0.4204   0.09808   0.09096  -0.0632   0.5184   1.0000
   7.250   0.4596   0.10060   0.09337  -0.0651   0.5159   1.0000
   7.500   0.5044   0.10387   0.09654  -0.0672   0.5146   1.0000
   7.750   0.4522   0.10444   0.09716  -0.0642   0.5018   1.0000
   8.000   0.4826   0.10662   0.09926  -0.0652   0.4986   1.0000
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