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

EPPLER 378 AIRFOIL (e378-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 378 AIRFOIL (e378-il)
Reynolds number: 200,000
Max Cl/Cd: 90.72 at α=7.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e378-il-200000-n5.txt
Download as CSV file: xf-e378-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 378 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.2535   0.10500   0.10148  -0.0042   0.8125   0.0122
  -7.750  -0.2461   0.10275   0.09906  -0.0044   0.7802   0.0125
  -7.500  -0.2384   0.10056   0.09672  -0.0047   0.7550   0.0127
  -7.250  -0.2280   0.09808   0.09430  -0.0064   0.7441   0.0130
  -7.000  -0.2179   0.09578   0.09204  -0.0078   0.7262   0.0133
  -6.500  -0.1957   0.09152   0.08761  -0.0112   0.7028   0.0137
  -6.250  -0.1809   0.08936   0.08540  -0.0140   0.6962   0.0138
  -6.000  -0.1652   0.08713   0.08306  -0.0169   0.6868   0.0139
  -5.750  -0.1480   0.08477   0.08066  -0.0198   0.6782   0.0139
  -5.500  -0.1296   0.08242   0.07821  -0.0230   0.6672   0.0140
  -5.250  -0.1080   0.07985   0.07560  -0.0270   0.6596   0.0140
  -5.000  -0.0853   0.07725   0.07294  -0.0310   0.6513   0.0141
  -4.750  -0.0612   0.07446   0.07011  -0.0350   0.6439   0.0141
  -4.500  -0.0368   0.07158   0.06718  -0.0386   0.6372   0.0141
  -4.250  -0.0231   0.06713   0.06274  -0.0390   0.6313   0.0144
  -4.000  -0.0031   0.06407   0.05962  -0.0409   0.6249   0.0147
  -3.750   0.0202   0.06123   0.05674  -0.0435   0.6182   0.0152
  -3.500   0.0470   0.05855   0.05398  -0.0469   0.6122   0.0159
  -3.250   0.0766   0.05582   0.05118  -0.0507   0.6055   0.0168
  -3.000   0.1235   0.05378   0.04900  -0.0583   0.5994   0.0202
  -2.750   0.1654   0.05063   0.04574  -0.0648   0.5936   0.0208
  -2.500   0.1843   0.04731   0.04240  -0.0653   0.5879   0.0215
  -2.250   0.2143   0.04480   0.03978  -0.0681   0.5827   0.0224
  -2.000   0.2487   0.04233   0.03723  -0.0715   0.5765   0.0234
  -1.750   0.2855   0.03982   0.03459  -0.0753   0.5713   0.0238
  -1.500   0.3263   0.03674   0.03132  -0.0796   0.5664   0.0181
  -1.250   0.3662   0.03388   0.02831  -0.0834   0.5609   0.0162
  -1.000   0.4087   0.03083   0.02503  -0.0873   0.5561   0.0153
  -0.750   0.4461   0.02849   0.02252  -0.0899   0.5512   0.0160
  -0.500   0.4864   0.02588   0.01967  -0.0927   0.5462   0.0198
  -0.250   0.5297   0.02190   0.01526  -0.0957   0.5421   0.0205
   0.000   0.5651   0.01936   0.01238  -0.0972   0.5374   0.0228
   0.250   0.6016   0.01552   0.00776  -0.0987   0.5314   0.0293
   0.500   0.6290   0.01521   0.00733  -0.0984   0.5268   0.0348
   0.750   0.6571   0.01440   0.00622  -0.0981   0.5216   0.0392
   1.000   0.6841   0.01398   0.00574  -0.0976   0.5165   0.0441
   1.250   0.7106   0.01379   0.00538  -0.0970   0.5122   0.0523
   1.500   0.7375   0.01371   0.00530  -0.0966   0.5074   0.0639
   1.750   0.7641   0.01366   0.00522  -0.0962   0.5026   0.0741
   2.000   0.7903   0.01371   0.00523  -0.0957   0.4984   0.0857
   2.250   0.8168   0.01367   0.00515  -0.0952   0.4938   0.0944
   2.500   0.8434   0.01365   0.00513  -0.0948   0.4891   0.1016
   2.750   0.8697   0.01369   0.00514  -0.0943   0.4848   0.1110
   3.000   0.8966   0.01371   0.00514  -0.0939   0.4803   0.1207
   3.250   0.9233   0.01376   0.00520  -0.0936   0.4754   0.1306
   3.500   0.9497   0.01384   0.00527  -0.0932   0.4710   0.1415
   3.750   0.9768   0.01392   0.00535  -0.0930   0.4665   0.1562
   4.000   1.0043   0.01398   0.00547  -0.0929   0.4613   0.1724
   4.250   1.0315   0.01407   0.00555  -0.0927   0.4566   0.1891
   4.500   1.0588   0.01417   0.00570  -0.0926   0.4514   0.2091
   4.750   1.0860   0.01427   0.00585  -0.0924   0.4454   0.2311
   5.000   1.1127   0.01440   0.00600  -0.0922   0.4398   0.2585
   5.250   1.1397   0.01451   0.00628  -0.0920   0.4325   0.2940
   5.500   1.1663   0.01460   0.00657  -0.0918   0.4262   0.4187
   6.000   1.2342   0.01386   0.00631  -0.0952   0.4104   1.0000
   6.250   1.2600   0.01406   0.00660  -0.0948   0.4014   1.0000
   6.500   1.2858   0.01427   0.00687  -0.0945   0.3922   1.0000
   6.750   1.3113   0.01450   0.00713  -0.0941   0.3824   1.0000
   7.000   1.3365   0.01475   0.00741  -0.0937   0.3710   1.0000
   7.250   1.3617   0.01501   0.00775  -0.0933   0.3582   1.0000
   7.500   1.3865   0.01532   0.00816  -0.0929   0.3441   1.0000
   7.750   1.4109   0.01567   0.00856  -0.0925   0.3272   1.0000
   8.000   1.4347   0.01610   0.00901  -0.0920   0.3046   1.0000
   8.250   1.4573   0.01667   0.00955  -0.0915   0.2768   1.0000
   8.500   1.4785   0.01746   0.01025  -0.0910   0.2428   1.0000
   8.750   1.4977   0.01852   0.01120  -0.0903   0.2060   1.0000
   9.000   1.5160   0.01966   0.01225  -0.0896   0.1747   1.0000
   9.250   1.5335   0.02085   0.01338  -0.0888   0.1482   1.0000
   9.500   1.5492   0.02218   0.01463  -0.0879   0.1222   1.0000
  10.000   1.5762   0.02502   0.01736  -0.0857   0.0802   1.0000
  10.250   1.5876   0.02646   0.01882  -0.0844   0.0663   1.0000
  10.500   1.5974   0.02793   0.02034  -0.0829   0.0564   1.0000
  10.750   1.6031   0.02957   0.02211  -0.0812   0.0488   1.0000
  11.000   1.6037   0.03133   0.02394  -0.0790   0.0425   1.0000
  11.250   1.6018   0.03355   0.02627  -0.0775   0.0377   1.0000
  11.500   1.6001   0.03610   0.02889  -0.0767   0.0322   1.0000
  11.750   1.5975   0.03891   0.03180  -0.0760   0.0276   1.0000
  12.000   1.5951   0.04184   0.03479  -0.0756   0.0230   1.0000
  12.250   1.5911   0.04521   0.03824  -0.0755   0.0196   1.0000
  12.500   1.5881   0.04851   0.04166  -0.0756   0.0165   1.0000
  12.750   1.5816   0.05235   0.04559  -0.0757   0.0147   1.0000
  13.000   1.5757   0.05625   0.04962  -0.0761   0.0129   1.0000
  13.250   1.5706   0.06005   0.05352  -0.0764   0.0104   1.0000
  13.500   1.5607   0.06461   0.05817  -0.0771   0.0092   1.0000
  13.750   1.5528   0.06904   0.06275  -0.0778   0.0083   1.0000
  14.000   1.5458   0.07335   0.06722  -0.0786   0.0069   1.0000
  14.250   1.5374   0.07815   0.07217  -0.0797   0.0063   1.0000
  14.500   1.5273   0.08335   0.07750  -0.0812   0.0059   1.0000
  14.750   1.5148   0.08917   0.08347  -0.0829   0.0055   1.0000
<< Back to EPPLER 378 AIRFOIL (e378-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 378 AIRFOIL (e378-il)