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

EPPLER 360 AIRFOIL (e360-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 360 AIRFOIL (e360-il)
Reynolds number: 50,000
Max Cl/Cd: 30.15 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e360-il-50000.txt
Download as CSV file: xf-e360-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 360 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.4802   0.13093   0.12376  -0.0039   1.0000   0.2336
 -10.750  -0.4608   0.12592   0.11875  -0.0032   1.0000   0.2417
 -10.500  -0.4835   0.12485   0.11783  -0.0055   1.0000   0.2507
 -10.250  -0.4678   0.12058   0.11356  -0.0046   1.0000   0.2649
 -10.000  -0.4499   0.11623   0.10923  -0.0038   1.0000   0.2773
  -9.750  -0.4427   0.11278   0.10582  -0.0035   1.0000   0.2902
  -9.500  -0.4379   0.10952   0.10262  -0.0031   1.0000   0.3068
  -9.250  -0.4328   0.10624   0.09940  -0.0027   1.0000   0.3224
  -9.000  -0.4277   0.10310   0.09630  -0.0021   1.0000   0.3388
  -8.750  -0.4224   0.09984   0.09309  -0.0016   1.0000   0.3544
  -8.500  -0.4171   0.09660   0.08992  -0.0010   1.0000   0.3700
  -8.250  -0.4121   0.09336   0.08674  -0.0003   1.0000   0.3858
  -8.000  -0.4071   0.09009   0.08353   0.0003   1.0000   0.4008
  -7.750  -0.4032   0.08683   0.08034   0.0010   1.0000   0.4152
  -7.250  -0.5886   0.06329   0.05653  -0.0257   1.0000   0.1536
  -7.000  -0.6115   0.05859   0.05096  -0.0226   1.0000   0.1268
  -6.750  -0.5971   0.05415   0.04665  -0.0214   1.0000   0.1230
  -6.500  -0.5943   0.05060   0.04283  -0.0188   1.0000   0.1180
  -6.250  -0.5948   0.04777   0.03926  -0.0149   1.0000   0.1111
  -6.000  -0.5865   0.04484   0.03615  -0.0123   1.0000   0.1100
  -5.750  -0.5784   0.04217   0.03321  -0.0094   1.0000   0.1089
  -5.500  -0.5689   0.03973   0.03048  -0.0066   1.0000   0.1076
  -5.250  -0.5579   0.03744   0.02787  -0.0038   1.0000   0.1069
  -5.000  -0.5453   0.03547   0.02556  -0.0012   1.0000   0.1079
  -4.750  -0.5318   0.03390   0.02357   0.0013   1.0000   0.1109
  -4.500  -0.5152   0.03194   0.02154   0.0030   1.0000   0.1163
  -4.250  -0.4962   0.03042   0.01988   0.0047   1.0000   0.1227
  -4.000  -0.4756   0.02891   0.01835   0.0061   1.0000   0.1346
  -3.750  -0.0241   0.02520   0.01646  -0.0398   1.0000   1.0000
  -3.500  -0.0152   0.02472   0.01589  -0.0385   1.0000   1.0000
  -3.250  -0.0087   0.02441   0.01553  -0.0369   1.0000   1.0000
  -3.000  -0.0068   0.02433   0.01542  -0.0345   1.0000   1.0000
  -2.750  -0.0142   0.02457   0.01567  -0.0308   1.0000   1.0000
  -2.500  -0.0296   0.02506   0.01614  -0.0261   1.0000   1.0000
  -2.250  -0.0440   0.02552   0.01656  -0.0217   1.0000   1.0000
  -2.000  -0.0553   0.02589   0.01687  -0.0177   1.0000   1.0000
  -1.750  -0.0639   0.02619   0.01712  -0.0141   1.0000   1.0000
  -1.500  -0.0704   0.02644   0.01728  -0.0107   1.0000   1.0000
  -1.250  -0.0753   0.02665   0.01742  -0.0075   1.0000   1.0000
  -1.000  -0.0789   0.02684   0.01752  -0.0044   1.0000   1.0000
  -0.750  -0.0339   0.02729   0.01780  -0.0098   0.9881   1.0000
  -0.500   0.0080   0.02774   0.01809  -0.0145   0.9761   1.0000
  -0.250   0.0442   0.02818   0.01841  -0.0180   0.9644   1.0000
   0.000   0.0787   0.02863   0.01876  -0.0211   0.9527   1.0000
   0.250   0.1133   0.02913   0.01917  -0.0240   0.9417   1.0000
   0.500   0.1500   0.02964   0.01961  -0.0272   0.9306   1.0000
   0.750   0.1712   0.03013   0.02005  -0.0276   0.9186   1.0000
   1.000   0.1959   0.03068   0.02056  -0.0285   0.9072   1.0000
   1.250   0.2309   0.03126   0.02112  -0.0310   0.8960   1.0000
   1.500   0.2569   0.03184   0.02168  -0.0319   0.8839   1.0000
   1.750   0.2729   0.03246   0.02229  -0.0310   0.8710   1.0000
   2.000   0.2945   0.03310   0.02293  -0.0310   0.8580   1.0000
   2.250   0.3225   0.03371   0.02357  -0.0318   0.8444   1.0000
   2.500   0.3544   0.03429   0.02418  -0.0330   0.8300   1.0000
   2.750   0.3876   0.03481   0.02475  -0.0342   0.8143   1.0000
   3.000   0.4198   0.03526   0.02527  -0.0350   0.7974   1.0000
   3.250   0.4558   0.03556   0.02567  -0.0360   0.7788   1.0000
   3.500   0.5069   0.03529   0.02552  -0.0384   0.7581   1.0000
   3.750   0.5681   0.03429   0.02470  -0.0409   0.7366   1.0000
   4.000   0.5806   0.03456   0.02503  -0.0376   0.7149   1.0000
   4.250   0.6244   0.03352   0.02410  -0.0371   0.6926   1.0000
   4.500   0.6527   0.03288   0.02358  -0.0347   0.6698   1.0000
   4.750   0.6805   0.03207   0.02285  -0.0320   0.6456   1.0000
   5.000   0.7048   0.03132   0.02217  -0.0289   0.6198   1.0000
   5.250   0.7363   0.02998   0.02088  -0.0261   0.5920   1.0000
   5.500   0.7585   0.02909   0.01999  -0.0225   0.5595   1.0000
   5.750   0.7780   0.02829   0.01914  -0.0187   0.5217   1.0000
   6.000   0.7986   0.02752   0.01822  -0.0151   0.4805   1.0000
   6.250   0.8172   0.02714   0.01756  -0.0117   0.4391   1.0000
   6.500   0.8304   0.02754   0.01777  -0.0084   0.4016   1.0000
   6.750   0.8463   0.02821   0.01821  -0.0058   0.3706   1.0000
   7.000   0.8640   0.02908   0.01886  -0.0037   0.3447   1.0000
   7.250   0.8801   0.03028   0.02000  -0.0016   0.3233   1.0000
   7.500   0.8992   0.03155   0.02117  -0.0001   0.3052   1.0000
   7.750   0.9178   0.03292   0.02251   0.0014   0.2893   1.0000
   8.000   0.9341   0.03457   0.02425   0.0030   0.2762   1.0000
   8.250   0.9502   0.03635   0.02614   0.0046   0.2648   1.0000
   8.500   0.9689   0.03810   0.02791   0.0058   0.2538   1.0000
   8.750   0.9828   0.04004   0.03006   0.0074   0.2444   1.0000
   9.000   0.9931   0.04242   0.03267   0.0093   0.2370   1.0000
   9.250   1.0086   0.04444   0.03479   0.0106   0.2283   1.0000
   9.500   1.0110   0.04722   0.03789   0.0130   0.2225   1.0000
   9.750   1.0070   0.05038   0.04141   0.0156   0.2183   1.0000
  10.000   1.0338   0.05261   0.04354   0.0156   0.2107   1.0000
  10.250   1.0134   0.05647   0.04785   0.0190   0.2088   1.0000
  10.500   0.9866   0.06084   0.05253   0.0220   0.2078   1.0000
  10.750   0.9512   0.06578   0.05766   0.0247   0.2082   1.0000
  11.000   0.9100   0.07167   0.06362   0.0259   0.2096   1.0000
  11.250   0.8724   0.07913   0.07110   0.0244   0.2109   1.0000
<< Back to EPPLER 360 AIRFOIL (e360-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 360 AIRFOIL (e360-il)