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=1,000,000 Ncrit=5


Details Polar file
Airfoil: EPPLER E850 AIRFOIL (e850-il)
Reynolds number: 1,000,000
Max Cl/Cd: 81.66 at α=0.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e850-il-1000000-n5.txt
Download as CSV file: xf-e850-il-1000000-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 =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -6.500  -0.3747   0.01487   0.01190  -0.0855   0.9693   0.0051
  -6.250  -0.3556   0.01305   0.00990  -0.0858   0.9663   0.0058
  -6.000  -0.3357   0.01134   0.00799  -0.0856   0.9625   0.0067
  -5.750  -0.3117   0.00953   0.00589  -0.0858   0.9601   0.0078
  -5.500  -0.2856   0.00797   0.00409  -0.0859   0.9584   0.0089
  -4.750  -0.2153   0.01248   0.00761  -0.0866   0.9563   0.0043
  -4.500  -0.1917   0.01036   0.00532  -0.0857   0.9532   0.0029
  -4.250  -0.1666   0.00922   0.00401  -0.0851   0.9499   0.0025
  -4.000  -0.1384   0.00842   0.00304  -0.0852   0.9475   0.0022
  -3.750  -0.1084   0.00794   0.00242  -0.0856   0.9458   0.0020
  -3.500  -0.0778   0.00764   0.00203  -0.0863   0.9443   0.0019
  -3.250  -0.0469   0.00744   0.00172  -0.0871   0.9430   0.0018
  -3.000  -0.0200   0.00730   0.00153  -0.0869   0.9400   0.0019
  -2.750   0.0080   0.00718   0.00137  -0.0870   0.9372   0.0021
  -2.500   0.0369   0.00706   0.00124  -0.0874   0.9348   0.0028
  -2.250   0.0666   0.00696   0.00117  -0.0879   0.9328   0.0037
  -2.000   0.0920   0.00523   0.00082  -0.0889   0.9303   0.4199
  -1.750   0.1184   0.00498   0.00083  -0.0888   0.9270   0.5036
  -1.500   0.1462   0.00492   0.00081  -0.0889   0.9233   0.5216
  -1.250   0.1755   0.00487   0.00078  -0.0893   0.9195   0.5355
  -1.000   0.2042   0.00483   0.00075  -0.0895   0.9143   0.5453
  -0.750   0.2320   0.00480   0.00072  -0.0895   0.9074   0.5524
  -0.500   0.2602   0.00478   0.00070  -0.0896   0.9006   0.5595
  -0.250   0.2882   0.00476   0.00069  -0.0897   0.8922   0.5664
   0.000   0.3154   0.00475   0.00068  -0.0895   0.8823   0.5734
   0.250   0.3426   0.00474   0.00068  -0.0894   0.8698   0.5809
   0.500   0.3694   0.00476   0.00067  -0.0891   0.8525   0.5881
   0.750   0.3944   0.00483   0.00067  -0.0884   0.8173   0.5960
   1.000   0.4102   0.00533   0.00071  -0.0857   0.7001   0.6033
   1.250   0.4223   0.00623   0.00095  -0.0825   0.5337   0.6110
   1.500   0.4470   0.00644   0.00113  -0.0819   0.4941   0.6196
   1.750   0.4659   0.00713   0.00137  -0.0804   0.3659   0.6282
   2.000   0.4854   0.00786   0.00162  -0.0790   0.2331   0.6377
   2.250   0.5086   0.00829   0.00183  -0.0783   0.1615   0.6479
   2.500   0.5321   0.00870   0.00206  -0.0776   0.0991   0.6586
   2.750   0.5560   0.00908   0.00230  -0.0770   0.0526   0.6702
   3.000   0.5809   0.00936   0.00253  -0.0766   0.0286   0.6828
   3.250   0.6056   0.00968   0.00285  -0.0761   0.0097   0.6960
   3.500   0.6303   0.01004   0.00327  -0.0754   0.0034   0.7103
   3.750   0.6543   0.01049   0.00387  -0.0745   0.0019   0.7263
   4.000   0.6791   0.01080   0.00427  -0.0740   0.0018   0.7435
   4.250   0.7042   0.01103   0.00459  -0.0736   0.0016   0.7629
   4.500   0.7288   0.01131   0.00499  -0.0730   0.0014   0.7848
   4.750   0.7526   0.01169   0.00551  -0.0722   0.0012   0.8094
   5.000   0.7757   0.01208   0.00605  -0.0713   0.0009   0.8391
   5.250   0.7968   0.01254   0.00671  -0.0699   0.0008   0.8807
   5.500   0.8190   0.01299   0.00741  -0.0686   0.0007   0.9946
   5.750   0.8418   0.01385   0.00838  -0.0676   0.0006   1.0000
   6.000   0.8644   0.01496   0.00964  -0.0666   0.0006   1.0000
   6.250   0.8867   0.01639   0.01134  -0.0654   0.0005   1.0000
   6.500   0.9071   0.01893   0.01426  -0.0637   0.0005   1.0000
   6.750   0.8963   0.03457   0.03133  -0.0528   0.0008   1.0000
   7.000   0.8992   0.04123   0.03842  -0.0481   0.0009   1.0000
   7.250   0.9017   0.04690   0.04440  -0.0444   0.0010   1.0000
   7.500   0.9019   0.05227   0.05002  -0.0411   0.0011   1.0000
   7.750   0.8999   0.05739   0.05535  -0.0383   0.0012   1.0000
   8.000   0.8939   0.06248   0.06062  -0.0358   0.0013   1.0000
   8.250   0.8871   0.06706   0.06533  -0.0338   0.0013   1.0000
   8.500   0.8739   0.07139   0.06977  -0.0315   0.0014   1.0000
   8.750   0.8558   0.07512   0.07357  -0.0291   0.0014   1.0000
   9.000   0.8390   0.07932   0.07784  -0.0286   0.0014   1.0000
<< Back to EPPLER E850 AIRFOIL (e850-il)

Polar data table (+)

Polar graphs


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