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

EPPLER 341 AIRFOIL (e341-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 341 AIRFOIL (e341-il)
Reynolds number: 500,000
Max Cl/Cd: 85.25 at α=10.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e341-il-500000.txt
Download as CSV file: xf-e341-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 341 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4744   0.10130   0.09929   0.0001   1.0000   0.0166
  -9.750  -0.4759   0.09630   0.09432  -0.0029   1.0000   0.0170
  -9.500  -0.4796   0.09061   0.08866  -0.0070   1.0000   0.0174
  -8.500  -0.4732   0.05219   0.04954  -0.0263   0.7714   0.0187
  -8.250  -0.4831   0.04917   0.04636  -0.0241   0.7487   0.0188
  -8.000  -0.5593   0.05633   0.05319  -0.0152   0.7838   0.0186
  -7.750  -0.5546   0.05401   0.05070  -0.0137   0.7559   0.0188
  -7.250  -0.5383   0.04962   0.04596  -0.0109   0.7119   0.0193
  -7.000  -0.5282   0.04719   0.04333  -0.0093   0.6936   0.0198
  -6.750  -0.5169   0.04454   0.04045  -0.0076   0.6774   0.0206
  -6.500  -0.4759   0.02452   0.02010  -0.0080   0.6444   0.0232
  -6.250  -0.4613   0.02266   0.01814  -0.0069   0.6313   0.0236
  -6.000  -0.4463   0.02077   0.01610  -0.0055   0.6187   0.0243
  -5.750  -0.4311   0.01867   0.01378  -0.0037   0.6075   0.0255
  -4.500  -0.3517   0.01830   0.01080   0.0101   0.5704   0.0144
  -4.250  -0.3238   0.01690   0.00925   0.0103   0.5600   0.0140
  -4.000  -0.2965   0.01587   0.00806   0.0108   0.5498   0.0138
  -3.750  -0.2700   0.01503   0.00712   0.0113   0.5400   0.0137
  -3.500  -0.2447   0.01438   0.00636   0.0119   0.5307   0.0139
  -3.250  -0.2198   0.01382   0.00574   0.0127   0.5212   0.0143
  -3.000  -0.1950   0.01341   0.00524   0.0134   0.5125   0.0147
  -2.750  -0.1719   0.01282   0.00463   0.0144   0.5041   0.0160
  -2.500  -0.1468   0.01253   0.00428   0.0151   0.4965   0.0176
  -2.250  -0.1220   0.01218   0.00391   0.0158   0.4887   0.0208
  -2.000  -0.0975   0.01187   0.00358   0.0166   0.4818   0.0306
  -1.750  -0.0758   0.01125   0.00330   0.0177   0.4747   0.1061
  -1.500  -0.0591   0.01038   0.00306   0.0195   0.4684   0.2835
  -1.250  -0.0749   0.00806   0.00269   0.0279   0.4646   0.7529
  -1.000  -0.0490   0.00898   0.00392   0.0303   0.4578   0.8838
  -0.750  -0.0211   0.01011   0.00494   0.0319   0.4518   0.9082
  -0.500   0.0403   0.01121   0.00594   0.0265   0.4443   0.9156
  -0.250   0.0951   0.01190   0.00648   0.0219   0.4372   0.9254
   0.000   0.1540   0.01245   0.00690   0.0163   0.4303   0.9330
   0.250   0.1937   0.01267   0.00704   0.0142   0.4245   0.9404
   0.500   0.2293   0.01271   0.00696   0.0125   0.4195   0.9424
   0.750   0.2610   0.01269   0.00690   0.0116   0.4150   0.9445
   1.000   0.2890   0.01269   0.00685   0.0115   0.4104   0.9474
   1.250   0.3085   0.01275   0.00685   0.0131   0.4063   0.9516
   1.500   0.3430   0.01274   0.00675   0.0115   0.4018   0.9526
   1.750   0.3765   0.01267   0.00666   0.0102   0.3979   0.9538
   2.000   0.4088   0.01264   0.00660   0.0092   0.3938   0.9552
   2.250   0.4389   0.01265   0.00655   0.0085   0.3897   0.9569
   2.500   0.4667   0.01273   0.00656   0.0083   0.3860   0.9589
   2.750   0.4859   0.01279   0.00664   0.0099   0.3832   0.9621
   3.000   0.5174   0.01273   0.00657   0.0090   0.3796   0.9632
   3.250   0.5501   0.01270   0.00651   0.0077   0.3758   0.9641
   3.500   0.5807   0.01274   0.00650   0.0069   0.3721   0.9650
   3.750   0.6108   0.01277   0.00652   0.0062   0.3688   0.9661
   4.000   0.6403   0.01276   0.00654   0.0056   0.3654   0.9675
   4.250   0.6675   0.01278   0.00657   0.0055   0.3620   0.9688
   4.500   0.6929   0.01284   0.00661   0.0058   0.3590   0.9702
   4.750   0.7129   0.01304   0.00676   0.0071   0.3557   0.9723
   5.000   0.7435   0.01301   0.00677   0.0062   0.3526   0.9728
   5.250   0.7731   0.01300   0.00681   0.0056   0.3491   0.9734
   5.500   0.8026   0.01301   0.00683   0.0049   0.3456   0.9741
   5.750   0.8322   0.01308   0.00689   0.0043   0.3422   0.9750
   6.000   0.8601   0.01323   0.00702   0.0039   0.3388   0.9758
   6.250   0.8866   0.01324   0.00711   0.0039   0.3354   0.9765
   6.500   0.9124   0.01328   0.00720   0.0040   0.3318   0.9772
   6.750   0.9381   0.01335   0.00728   0.0041   0.3281   0.9782
   7.000   0.9627   0.01357   0.00745   0.0044   0.3241   0.9793
   7.250   0.9857   0.01361   0.00759   0.0050   0.3211   0.9800
   7.500   1.0071   0.01368   0.00773   0.0060   0.3172   0.9808
   7.750   1.0339   0.01373   0.00781   0.0058   0.3130   0.9813
   8.000   1.0623   0.01391   0.00797   0.0053   0.3086   0.9820
   8.250   1.0895   0.01392   0.00810   0.0050   0.3044   0.9826
   8.500   1.1148   0.01398   0.00821   0.0051   0.2996   0.9830
   8.750   1.1388   0.01415   0.00837   0.0054   0.2946   0.9836
   9.000   1.1633   0.01423   0.00856   0.0056   0.2899   0.9841
   9.250   1.1871   0.01433   0.00872   0.0060   0.2845   0.9847
   9.500   1.2102   0.01454   0.00892   0.0064   0.2788   0.9856
   9.750   1.2340   0.01463   0.00914   0.0068   0.2729   0.9865
  10.000   1.2541   0.01483   0.00936   0.0078   0.2667   0.9872
  10.250   1.2737   0.01501   0.00961   0.0089   0.2599   0.9880
  10.500   1.2940   0.01525   0.00987   0.0098   0.2526   0.9887
  10.750   1.3180   0.01546   0.01016   0.0099   0.2442   0.9892
  11.000   1.3399   0.01578   0.01049   0.0103   0.2346   0.9899
  11.250   1.3610   0.01614   0.01088   0.0107   0.2241   0.9907
  11.500   1.3826   0.01656   0.01133   0.0110   0.2125   0.9919
  11.750   1.3998   0.01707   0.01186   0.0121   0.2006   0.9930
  12.000   1.4135   0.01768   0.01247   0.0136   0.1883   0.9943
  12.250   1.4270   0.01839   0.01317   0.0150   0.1758   0.9955
  12.500   1.4416   0.01924   0.01401   0.0159   0.1619   0.9967
  12.750   1.4517   0.02021   0.01498   0.0172   0.1490   0.9983
  13.000   1.4480   0.02121   0.01599   0.0209   0.1399   1.0000
  13.250   1.4286   0.02210   0.01694   0.0274   0.1351   1.0000
  13.500   1.4057   0.02340   0.01828   0.0333   0.1309   1.0000
  13.750   1.3835   0.02496   0.01990   0.0384   0.1272   1.0000
  14.000   1.3686   0.02682   0.02181   0.0416   0.1218   1.0000
  14.250   1.3611   0.02970   0.02473   0.0417   0.1147   1.0000
  14.500   1.3569   0.03286   0.02793   0.0407   0.1070   1.0000
  14.750   1.3503   0.03646   0.03159   0.0395   0.1004   1.0000
  15.000   1.3406   0.04047   0.03564   0.0382   0.0943   1.0000
  15.250   1.3309   0.04448   0.03971   0.0369   0.0891   1.0000
  15.500   1.3193   0.04872   0.04401   0.0356   0.0842   1.0000
  15.750   1.3046   0.05331   0.04865   0.0343   0.0801   1.0000
  16.000   1.2955   0.05739   0.05280   0.0330   0.0758   1.0000
  16.250   1.2817   0.06212   0.05757   0.0315   0.0720   1.0000
  16.500   1.2711   0.06657   0.06209   0.0300   0.0684   1.0000
  16.750   1.2620   0.07088   0.06646   0.0285   0.0647   1.0000
  17.000   1.2491   0.07574   0.07135   0.0268   0.0613   1.0000
  17.250   1.2415   0.07997   0.07564   0.0253   0.0581   1.0000
  17.500   1.2331   0.08433   0.08006   0.0238   0.0548   1.0000
  17.750   1.2210   0.08925   0.08502   0.0220   0.0519   1.0000
<< Back to EPPLER 341 AIRFOIL (e341-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 341 AIRFOIL (e341-il)