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

EPPLER E851 AIRFOIL (e851-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: EPPLER E851 AIRFOIL (e851-il)
Reynolds number: 50,000
Max Cl/Cd: 31.02 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e851-il-50000-n5.txt
Download as CSV file: xf-e851-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER E851 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4818   0.08641   0.07960  -0.0473   1.0000   0.0295
  -8.250  -0.4901   0.08281   0.07609  -0.0475   1.0000   0.0292
  -8.000  -0.5019   0.07952   0.07289  -0.0475   1.0000   0.0289
  -7.750  -0.5131   0.07597   0.06941  -0.0480   1.0000   0.0287
  -7.500  -0.5229   0.07242   0.06586  -0.0481   1.0000   0.0284
  -7.250  -0.5312   0.06893   0.06234  -0.0480   1.0000   0.0282
  -7.000  -0.5378   0.06522   0.05855  -0.0479   1.0000   0.0279
  -6.750  -0.5411   0.06147   0.05465  -0.0477   1.0000   0.0277
  -6.500  -0.5404   0.05777   0.05075  -0.0476   1.0000   0.0274
  -6.250  -0.5362   0.05384   0.04654  -0.0476   1.0000   0.0271
  -6.000  -0.5279   0.04991   0.04225  -0.0475   1.0000   0.0270
  -5.750  -0.5156   0.04616   0.03808  -0.0474   1.0000   0.0269
  -5.500  -0.4995   0.04246   0.03388  -0.0473   1.0000   0.0270
  -5.250  -0.4803   0.03906   0.02994  -0.0470   1.0000   0.0274
  -5.000  -0.4587   0.03602   0.02635  -0.0464   1.0000   0.0280
  -4.750  -0.4357   0.03343   0.02323  -0.0455   1.0000   0.0291
  -4.500  -0.4138   0.03101   0.02057  -0.0447   1.0000   0.0321
  -4.250  -0.3914   0.02964   0.01888  -0.0438   1.0000   0.0410
  -4.000  -0.3698   0.02770   0.01689  -0.0424   1.0000   0.0478
  -3.750  -0.3474   0.02597   0.01479  -0.0411   1.0000   0.0593
  -3.500  -0.3241   0.02447   0.01327  -0.0408   1.0000   0.0869
  -3.250  -0.2960   0.02217   0.01140  -0.0420   1.0000   0.1660
  -3.000  -0.2770   0.02053   0.01148  -0.0409   1.0000   0.5263
  -2.750  -0.2627   0.02067   0.01170  -0.0372   1.0000   0.6182
  -2.500  -0.2519   0.02083   0.01190  -0.0326   1.0000   0.6834
  -2.250  -0.2396   0.02086   0.01186  -0.0286   1.0000   0.7299
  -2.000  -0.2230   0.02075   0.01160  -0.0260   1.0000   0.7581
  -1.750  -0.2029   0.02062   0.01125  -0.0247   1.0000   0.7781
  -1.500  -0.1821   0.02049   0.01094  -0.0236   1.0000   0.7982
  -1.250  -0.1617   0.02037   0.01054  -0.0224   1.0000   0.8197
  -1.000  -0.1414   0.02024   0.01030  -0.0213   1.0000   0.8439
  -0.750  -0.1195   0.02012   0.01010  -0.0205   1.0000   0.8751
  -0.500  -0.0898   0.01998   0.00989  -0.0216   1.0000   0.9225
  -0.250  -0.0526   0.01992   0.00967  -0.0250   0.9960   1.0000
   0.000  -0.0157   0.02030   0.00981  -0.0283   0.9918   1.0000
   0.250   0.0194   0.02066   0.00990  -0.0310   0.9870   1.0000
   0.500   0.0555   0.02114   0.01021  -0.0338   0.9825   1.0000
   0.750   0.0888   0.02154   0.01048  -0.0359   0.9773   1.0000
   1.000   0.1226   0.02199   0.01083  -0.0381   0.9718   1.0000
   1.250   0.1566   0.02245   0.01123  -0.0402   0.9664   1.0000
   1.500   0.1887   0.02287   0.01163  -0.0420   0.9599   1.0000
   1.750   0.2232   0.02336   0.01211  -0.0441   0.9540   1.0000
   2.000   0.2550   0.02378   0.01255  -0.0457   0.9467   1.0000
   2.250   0.2883   0.02424   0.01307  -0.0475   0.9397   1.0000
   2.500   0.3227   0.02470   0.01362  -0.0495   0.9321   1.0000
   2.750   0.3539   0.02511   0.01427  -0.0508   0.9232   1.0000
   3.000   0.3897   0.02555   0.01486  -0.0529   0.9148   1.0000
   3.250   0.4253   0.02594   0.01545  -0.0548   0.9052   1.0000
   3.500   0.4582   0.02626   0.01599  -0.0561   0.8937   1.0000
   3.750   0.4934   0.02652   0.01652  -0.0576   0.8812   1.0000
   4.000   0.5320   0.02663   0.01712  -0.0594   0.8669   1.0000
   4.250   0.5708   0.02648   0.01737  -0.0608   0.8484   1.0000
   4.500   0.6172   0.02548   0.01687  -0.0619   0.8180   1.0000
   4.750   0.6669   0.02302   0.01496  -0.0608   0.7598   1.0000
   5.000   0.6834   0.02203   0.01425  -0.0560   0.6762   1.0000
   5.250   0.7199   0.02375   0.01307  -0.0534   0.1931   1.0000
   5.500   0.7314   0.02626   0.01480  -0.0513   0.1071   1.0000
   5.750   0.7505   0.02835   0.01675  -0.0499   0.0760   1.0000
   6.000   0.7796   0.03066   0.01920  -0.0495   0.0575   1.0000
   6.250   0.8148   0.03322   0.02216  -0.0499   0.0408   1.0000
   6.500   0.8531   0.03678   0.02592  -0.0508   0.0346   1.0000
   6.750   0.8774   0.03905   0.02858  -0.0500   0.0277   1.0000
   7.000   0.8983   0.04260   0.03238  -0.0492   0.0245   1.0000
   7.250   0.9183   0.04623   0.03661  -0.0475   0.0237   1.0000
   7.500   0.9323   0.05000   0.04094  -0.0454   0.0232   1.0000
   7.750   0.9410   0.05391   0.04536  -0.0429   0.0229   1.0000
   8.000   0.9454   0.05783   0.04973  -0.0403   0.0227   1.0000
   8.250   0.9458   0.06178   0.05409  -0.0376   0.0227   1.0000
   8.500   0.9424   0.06570   0.05835  -0.0350   0.0227   1.0000
   8.750   0.9358   0.06955   0.06248  -0.0326   0.0228   1.0000
   9.000   0.9249   0.07325   0.06641  -0.0301   0.0229   1.0000
   9.250   0.9112   0.07689   0.07022  -0.0279   0.0230   1.0000
   9.500   0.8959   0.08083   0.07431  -0.0266   0.0231   1.0000
   9.750   0.8803   0.08512   0.07874  -0.0264   0.0233   1.0000
  10.000   0.8645   0.08997   0.08368  -0.0275   0.0234   1.0000
  10.250   0.8504   0.09527   0.08905  -0.0297   0.0236   1.0000
<< Back to EPPLER E851 AIRFOIL (e851-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER E851 AIRFOIL (e851-il)