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

EPPLER E850 AIRFOIL (e850-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: EPPLER E850 AIRFOIL (e850-il)
Reynolds number: 500,000
Max Cl/Cd: 77.62 at α=1.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e850-il-500000-n5.txt
Download as CSV file: xf-e850-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER E850 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5157   0.08839   0.08618  -0.0290   1.0000   0.0087
  -8.500  -0.5204   0.08486   0.08268  -0.0296   1.0000   0.0090
  -8.250  -0.5250   0.08067   0.07852  -0.0313   0.9996   0.0090
  -8.000  -0.5192   0.07431   0.07218  -0.0384   0.9971   0.0090
  -7.750  -0.5141   0.06541   0.06323  -0.0524   0.9926   0.0089
  -7.500  -0.5056   0.05864   0.05631  -0.0603   0.9875   0.0091
  -7.250  -0.4919   0.05275   0.05023  -0.0660   0.9839   0.0093
  -5.500  -0.3571   0.02116   0.01624  -0.0747   0.9627   0.0057
  -5.250  -0.3275   0.01968   0.01443  -0.0743   0.9610   0.0042
  -5.000  -0.3003   0.01729   0.01177  -0.0744   0.9596   0.0036
  -4.750  -0.2720   0.01511   0.00931  -0.0745   0.9584   0.0033
  -4.500  -0.2430   0.01346   0.00746  -0.0748   0.9574   0.0031
  -4.250  -0.2133   0.01222   0.00606  -0.0753   0.9565   0.0030
  -4.000  -0.1827   0.01121   0.00490  -0.0761   0.9558   0.0029
  -3.750  -0.1613   0.01060   0.00415  -0.0747   0.9516   0.0028
  -3.500  -0.1329   0.01013   0.00352  -0.0749   0.9496   0.0028
  -3.250  -0.1030   0.00981   0.00307  -0.0754   0.9481   0.0029
  -3.000  -0.0722   0.00959   0.00271  -0.0761   0.9469   0.0029
  -2.750  -0.0408   0.00942   0.00248  -0.0770   0.9459   0.0034
  -2.500  -0.0091   0.00919   0.00228  -0.0780   0.9450   0.0213
  -2.250   0.0206   0.00801   0.00196  -0.0795   0.9441   0.2569
  -2.000   0.0492   0.00701   0.00201  -0.0806   0.9432   0.5418
  -1.750   0.0742   0.00698   0.00201  -0.0800   0.9398   0.5595
  -1.500   0.1023   0.00693   0.00196  -0.0802   0.9371   0.5710
  -1.250   0.1324   0.00687   0.00191  -0.0807   0.9351   0.5793
  -1.000   0.1639   0.00681   0.00185  -0.0816   0.9333   0.5874
  -0.750   0.1962   0.00673   0.00180  -0.0827   0.9317   0.5957
  -0.500   0.2298   0.00664   0.00174  -0.0840   0.9300   0.6038
   0.000   0.2866   0.00650   0.00167  -0.0843   0.9198   0.6209
   0.250   0.3181   0.00641   0.00163  -0.0850   0.9149   0.6298
   0.500   0.3470   0.00633   0.00159  -0.0852   0.9070   0.6393
   0.750   0.3762   0.00624   0.00157  -0.0854   0.8974   0.6492
   1.000   0.4055   0.00616   0.00153  -0.0856   0.8841   0.6593
   1.250   0.4343   0.00608   0.00148  -0.0856   0.8612   0.6701
   1.500   0.4619   0.00606   0.00152  -0.0852   0.8223   0.6816
   1.750   0.4859   0.00626   0.00150  -0.0841   0.7442   0.6938
   2.000   0.4967   0.00711   0.00169  -0.0803   0.5879   0.7064
   2.250   0.5073   0.00821   0.00206  -0.0771   0.4023   0.7200
   2.500   0.5238   0.00916   0.00243  -0.0752   0.2436   0.7351
   2.750   0.5447   0.00977   0.00275  -0.0741   0.1510   0.7523
   3.000   0.5663   0.01033   0.00309  -0.0731   0.0778   0.7711
   3.250   0.5889   0.01077   0.00344  -0.0721   0.0367   0.7923
   3.500   0.6113   0.01118   0.00389  -0.0711   0.0122   0.8169
   3.750   0.6313   0.01195   0.00492  -0.0690   0.0055   0.8466
   4.000   0.6501   0.01254   0.00575  -0.0668   0.0052   0.8884
   4.250   0.6773   0.01284   0.00621  -0.0667   0.0040   0.9913
   4.500   0.7032   0.01299   0.00632  -0.0667   0.0026   1.0000
   4.750   0.7248   0.01402   0.00746  -0.0654   0.0021   1.0000
   5.000   0.7466   0.01525   0.00883  -0.0642   0.0018   1.0000
   5.250   0.7704   0.01621   0.00992  -0.0634   0.0016   1.0000
   5.500   0.7940   0.01776   0.01168  -0.0623   0.0014   1.0000
   5.750   0.8175   0.02010   0.01436  -0.0610   0.0012   1.0000
   6.000   0.8391   0.02325   0.01795  -0.0592   0.0012   1.0000
   6.250   0.8566   0.02723   0.02244  -0.0565   0.0012   1.0000
   6.500   0.8690   0.03229   0.02802  -0.0529   0.0012   1.0000
   6.750   0.8784   0.03743   0.03359  -0.0493   0.0012   1.0000
   7.000   0.8853   0.04246   0.03898  -0.0459   0.0013   1.0000
   7.250   0.8898   0.04744   0.04426  -0.0427   0.0013   1.0000
   7.500   0.8913   0.05242   0.04951  -0.0397   0.0014   1.0000
   7.750   0.8902   0.05729   0.05461  -0.0370   0.0014   1.0000
   8.000   0.8855   0.06213   0.05964  -0.0346   0.0015   1.0000
   8.250   0.8786   0.06663   0.06429  -0.0326   0.0015   1.0000
   8.500   0.8659   0.07081   0.06859  -0.0303   0.0015   1.0000
   8.750   0.8491   0.07472   0.07259  -0.0283   0.0015   1.0000
   9.000   0.8326   0.07916   0.07710  -0.0282   0.0015   1.0000
   9.250   0.8168   0.08458   0.08257  -0.0306   0.0015   1.0000
<< Back to EPPLER E850 AIRFOIL (e850-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER E850 AIRFOIL (e850-il)