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

EPPLER 856 AIRFOIL (e856-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 856 AIRFOIL (e856-il)
Reynolds number: 100,000
Max Cl/Cd: 46.34 at α=7.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e856-il-100000-n5.txt
Download as CSV file: xf-e856-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 856 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.750  -0.4007   0.10303   0.09796  -0.0604   1.0000   0.0256
 -12.500  -0.4644   0.08616   0.08096  -0.0692   1.0000   0.0245
 -12.250  -0.5259   0.07341   0.06791  -0.0761   1.0000   0.0233
 -12.000  -0.5591   0.06685   0.06117  -0.0786   1.0000   0.0230
 -11.750  -0.5901   0.06134   0.05541  -0.0798   1.0000   0.0226
 -11.500  -0.6027   0.05881   0.05288  -0.0793   1.0000   0.0230
 -11.250  -0.6225   0.05588   0.04987  -0.0783   1.0000   0.0230
 -11.000  -0.6308   0.05297   0.04686  -0.0792   0.9976   0.0232
 -10.750  -0.6181   0.04948   0.04316  -0.0845   0.9894   0.0240
 -10.500  -0.6085   0.04614   0.03955  -0.0887   0.9793   0.0250
 -10.250  -0.6005   0.04312   0.03621  -0.0909   0.9668   0.0258
 -10.000  -0.5858   0.04054   0.03339  -0.0921   0.9548   0.0267
  -9.750  -0.5671   0.03863   0.03142  -0.0933   0.9433   0.0277
  -9.500  -0.5450   0.03669   0.02934  -0.0947   0.9330   0.0291
  -9.250  -0.5229   0.03479   0.02718  -0.0956   0.9225   0.0313
  -9.000  -0.5026   0.03330   0.02569  -0.0964   0.9105   0.0336
  -8.750  -0.4778   0.03168   0.02386  -0.0971   0.9009   0.0365
  -8.500  -0.4521   0.03016   0.02236  -0.0983   0.8914   0.0395
  -8.250  -0.4283   0.02884   0.02092  -0.0989   0.8798   0.0438
  -8.000  -0.3974   0.02751   0.01947  -0.1006   0.8712   0.0496
  -7.750  -0.3698   0.02628   0.01822  -0.1018   0.8599   0.0563
  -7.500  -0.3404   0.02514   0.01702  -0.1033   0.8486   0.0649
  -7.250  -0.3032   0.02397   0.01576  -0.1061   0.8400   0.0770
  -7.000  -0.2751   0.02301   0.01476  -0.1071   0.8270   0.0901
  -6.750  -0.2458   0.02199   0.01375  -0.1084   0.8147   0.1066
  -6.500  -0.2139   0.02102   0.01274  -0.1102   0.8035   0.1293
  -6.250  -0.1853   0.02010   0.01184  -0.1113   0.7911   0.1578
  -6.000  -0.1595   0.01922   0.01106  -0.1119   0.7783   0.1977
  -5.750  -0.1325   0.01842   0.01048  -0.1126   0.7666   0.2600
  -5.500  -0.1027   0.01809   0.01029  -0.1134   0.7556   0.3248
  -5.250  -0.0755   0.01808   0.01026  -0.1132   0.7435   0.3646
  -5.000  -0.0455   0.01814   0.01015  -0.1136   0.7327   0.3936
  -4.750  -0.0156   0.01825   0.01009  -0.1138   0.7221   0.4150
  -4.500   0.0114   0.01835   0.01005  -0.1135   0.7113   0.4319
  -4.250   0.0415   0.01845   0.00995  -0.1138   0.7023   0.4471
  -4.000   0.0676   0.01852   0.00986  -0.1135   0.6921   0.4606
  -3.750   0.0961   0.01864   0.00983  -0.1134   0.6835   0.4734
  -3.500   0.1222   0.01876   0.00987  -0.1129   0.6740   0.4845
  -3.250   0.1499   0.01883   0.00977  -0.1128   0.6659   0.4965
  -3.000   0.1760   0.01891   0.00977  -0.1123   0.6573   0.5058
  -2.750   0.2045   0.01891   0.00958  -0.1125   0.6503   0.5140
  -2.500   0.2299   0.01892   0.00955  -0.1120   0.6421   0.5200
  -2.250   0.2588   0.01890   0.00932  -0.1123   0.6355   0.5273
  -2.000   0.2839   0.01890   0.00926  -0.1119   0.6277   0.5327
  -1.750   0.3110   0.01891   0.00918  -0.1117   0.6210   0.5382
  -1.500   0.3386   0.01893   0.00905  -0.1118   0.6149   0.5448
  -1.250   0.3643   0.01895   0.00902  -0.1115   0.6082   0.5500
  -1.000   0.3920   0.01899   0.00896  -0.1115   0.6026   0.5554
  -0.750   0.4182   0.01904   0.00893  -0.1113   0.5965   0.5618
  -0.500   0.4440   0.01909   0.00893  -0.1111   0.5903   0.5673
  -0.250   0.4717   0.01914   0.00891  -0.1111   0.5854   0.5727
   0.000   0.4978   0.01923   0.00894  -0.1109   0.5802   0.5793
   0.250   0.5231   0.01932   0.00902  -0.1106   0.5747   0.5851
   0.500   0.5501   0.01940   0.00906  -0.1105   0.5699   0.5911
   0.750   0.5776   0.01951   0.00908  -0.1106   0.5654   0.5981
   1.000   0.6012   0.01964   0.00926  -0.1099   0.5599   0.6040
   1.250   0.6274   0.01976   0.00937  -0.1097   0.5553   0.6110
   1.500   0.6556   0.01989   0.00942  -0.1099   0.5515   0.6184
   1.750   0.6801   0.02005   0.00962  -0.1094   0.5472   0.6250
   2.000   0.7039   0.02023   0.00982  -0.1089   0.5423   0.6335
   2.250   0.7292   0.02037   0.00999  -0.1085   0.5380   0.6409
   2.500   0.7572   0.02052   0.01010  -0.1087   0.5345   0.6500
   2.750   0.7809   0.02072   0.01037  -0.1081   0.5306   0.6583
   3.000   0.8037   0.02097   0.01068  -0.1074   0.5263   0.6686
   3.250   0.8282   0.02116   0.01093  -0.1069   0.5222   0.6784
   3.500   0.8553   0.02133   0.01111  -0.1069   0.5188   0.6900
   3.750   0.8815   0.02154   0.01133  -0.1067   0.5154   0.7024
   4.000   0.9007   0.02183   0.01179  -0.1054   0.5111   0.7154
   4.250   0.9228   0.02209   0.01215  -0.1045   0.5074   0.7306
   4.500   0.9470   0.02229   0.01243  -0.1039   0.5040   0.7483
   4.750   0.9734   0.02244   0.01264  -0.1037   0.5010   0.7692
   5.000   0.9929   0.02270   0.01304  -0.1023   0.4974   0.7945
   5.250   1.0081   0.02297   0.01353  -0.1001   0.4934   0.8299
   5.500   1.0305   0.02315   0.01393  -0.0991   0.4899   0.8965
   5.750   1.0626   0.02341   0.01420  -0.1004   0.4866   1.0000
   6.000   1.0928   0.02371   0.01444  -0.1012   0.4838   1.0000
   6.250   1.1098   0.02428   0.01510  -0.0998   0.4799   1.0000
   6.500   1.1266   0.02486   0.01576  -0.0984   0.4760   1.0000
   6.750   1.1477   0.02533   0.01628  -0.0977   0.4725   1.0000
   7.000   1.1735   0.02571   0.01667  -0.0976   0.4695   1.0000
   7.250   1.2045   0.02599   0.01692  -0.0985   0.4668   1.0000
   7.500   1.2111   0.02681   0.01790  -0.0954   0.4625   1.0000
   7.750   1.2234   0.02749   0.01870  -0.0933   0.4585   1.0000
   8.000   1.2430   0.02798   0.01925  -0.0923   0.4549   1.0000
   8.250   1.2707   0.02829   0.01957  -0.0925   0.4519   1.0000
   8.500   1.2905   0.02883   0.02017  -0.0916   0.4486   1.0000
   8.750   1.2840   0.02986   0.02139  -0.0866   0.4437   1.0000
   9.000   1.2955   0.03050   0.02212  -0.0844   0.4394   1.0000
   9.250   1.3218   0.03074   0.02239  -0.0843   0.4359   1.0000
   9.500   1.3393   0.03123   0.02294  -0.0830   0.4319   1.0000
   9.750   1.3234   0.03273   0.02464  -0.0772   0.4263   1.0000
  10.000   1.3375   0.03324   0.02522  -0.0756   0.4216   1.0000
  10.250   1.3770   0.03286   0.02482  -0.0771   0.4177   1.0000
  10.500   1.3431   0.03528   0.02747  -0.0699   0.4113   1.0000
  10.750   1.3485   0.03618   0.02846  -0.0675   0.4057   1.0000
  11.000   1.3930   0.03520   0.02746  -0.0691   0.4011   1.0000
  11.250   1.3356   0.03960   0.03212  -0.0614   0.3936   1.0000
  11.500   1.3556   0.03972   0.03229  -0.0606   0.3879   1.0000
  11.750   1.3400   0.04237   0.03507  -0.0577   0.3814   1.0000
  12.000   1.3138   0.04620   0.03903  -0.0548   0.3735   1.0000
  12.250   1.3573   0.04439   0.03723  -0.0551   0.3688   1.0000
  12.500   1.2769   0.05383   0.04686  -0.0517   0.3571   1.0000
  12.750   1.2181   0.06289   0.05602  -0.0510   0.3435   1.0000
  13.000   1.2613   0.06020   0.05340  -0.0503   0.3404   1.0000
  13.500   1.1935   0.07404   0.06742  -0.0507   0.3149   1.0000
  13.750   1.2221   0.07292   0.06637  -0.0500   0.3103   1.0000
  14.250   1.2373   0.07638   0.07000  -0.0496   0.2936   1.0000
  14.500   1.2171   0.08210   0.07579  -0.0503   0.2807   1.0000
  14.750   1.2120   0.08583   0.07961  -0.0508   0.2698   1.0000
  15.000   1.2430   0.08420   0.07801  -0.0499   0.2632   1.0000
  15.250   1.2330   0.08872   0.08260  -0.0506   0.2508   1.0000
  15.500   1.2344   0.09157   0.08551  -0.0509   0.2395   1.0000
  15.750   1.2465   0.09276   0.08671  -0.0508   0.2290   1.0000
  16.000   1.2606   0.09359   0.08750  -0.0506   0.2180   1.0000
  16.250   1.2603   0.09674   0.09068  -0.0511   0.2061   1.0000
  16.500   1.2602   0.09990   0.09384  -0.0517   0.1947   1.0000
  16.750   1.2627   0.10265   0.09658  -0.0521   0.1837   1.0000
  17.000   1.2663   0.10518   0.09906  -0.0526   0.1732   1.0000
  17.250   1.2646   0.10868   0.10257  -0.0534   0.1631   1.0000
  17.500   1.2631   0.11220   0.10612  -0.0544   0.1537   1.0000
  17.750   1.2647   0.11516   0.10903  -0.0552   0.1450   1.0000
<< Back to EPPLER 856 AIRFOIL (e856-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 856 AIRFOIL (e856-il)