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

Eiffel 10 (Wright) - 1903 Wright Flyer airfoil (eiffel10-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: Eiffel 10 (Wright) - 1903 Wright Flyer airfoil (eiffel10-il)
Reynolds number: 200,000
Max Cl/Cd: 28.74 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-eiffel10-il-200000.txt
Download as CSV file: xf-eiffel10-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: Eiffel 10 (Wright) - 1903 Wright Flyer 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.500  -0.3732   0.13029   0.12646  -0.0026   1.0000   0.0183
 -11.250  -0.3683   0.12766   0.12384  -0.0037   1.0000   0.0183
  -8.750  -0.4196   0.11180   0.10789  -0.0010   1.0000   0.0184
  -8.500  -0.4108   0.10839   0.10451  -0.0015   1.0000   0.0185
  -8.250  -0.4022   0.10543   0.10158  -0.0025   1.0000   0.0187
  -8.000  -0.3937   0.10264   0.09883  -0.0037   1.0000   0.0188
  -7.750  -0.3851   0.09996   0.09618  -0.0049   1.0000   0.0190
  -7.500  -0.3752   0.09725   0.09351  -0.0066   1.0000   0.0192
  -7.250  -0.3629   0.09442   0.09072  -0.0088   1.0000   0.0194
  -7.000  -0.3496   0.09157   0.08790  -0.0113   1.0000   0.0196
  -6.750  -0.3351   0.08869   0.08506  -0.0140   1.0000   0.0199
  -6.500  -0.3194   0.08581   0.08221  -0.0169   1.0000   0.0202
  -6.250  -0.3023   0.08292   0.07936  -0.0201   1.0000   0.0205
  -6.000  -0.2830   0.08010   0.07657  -0.0239   1.0000   0.0209
  -5.750  -0.2670   0.06807   0.06504  -0.0173   0.9998   0.0212
  -5.500  -0.2338   0.06184   0.05882  -0.0224   0.9948   0.0218
  -5.250  -0.1871   0.05580   0.05275  -0.0322   0.9867   0.0226
  -5.000  -0.1466   0.05088   0.04781  -0.0408   0.9795   0.0238
  -4.750  -0.1008   0.04555   0.04244  -0.0502   0.9623   0.0247
  -4.500  -0.1133   0.06187   0.05842  -0.0555   0.9865   0.0243
  -4.250  -0.0810   0.05742   0.05401  -0.0590   0.9679   0.0250
  -4.000  -0.0396   0.05398   0.05054  -0.0652   0.9474   0.0260
  -3.750   0.0007   0.05094   0.04738  -0.0706   0.9134   0.0272
  -3.250   0.0613   0.04594   0.04200  -0.0755   0.8234   0.0283
  -3.000   0.0834   0.04411   0.03969  -0.0754   0.7206   0.0293
  -2.750   0.1050   0.04499   0.03801  -0.0768   0.0574   0.0304
  -2.500   0.1570   0.04280   0.03557  -0.0830   0.0500   0.0318
  -2.250   0.1789   0.04042   0.03326  -0.0836   0.0475   0.0328
  -2.000   0.2145   0.03848   0.03126  -0.0863   0.0462   0.0349
  -1.750   0.2557   0.03642   0.02907  -0.0897   0.0456   0.0367
  -1.500   0.2858   0.03474   0.02741  -0.0911   0.0454   0.0388
  -1.250   0.3271   0.03307   0.02556  -0.0938   0.0455   0.0418
  -1.000   0.3563   0.03169   0.02421  -0.0949   0.0460   0.0447
  -0.750   0.3926   0.03035   0.02272  -0.0966   0.0468   0.0486
  -0.500   0.4291   0.02965   0.02175  -0.0978   0.0478   0.0548
  -0.250   0.4558   0.02873   0.02084  -0.0984   0.0493   0.0592
   0.000   0.4864   0.02842   0.02029  -0.0990   0.0507   0.0668
   0.250   0.5190   0.02733   0.01918  -0.0994   0.0537   0.0774
   0.500   0.5488   0.02717   0.01891  -0.0995   0.0566   0.0911
   0.750   0.5793   0.02734   0.01885  -0.0996   0.0575   0.1144
   1.000   0.6092   0.02700   0.01856  -0.0994   0.0638   0.1449
   1.250   0.6384   0.02711   0.01884  -0.0989   0.0791   0.1818
   3.500   0.9241   0.03215   0.02365  -0.0901   0.1145   0.0455
   3.750   0.9476   0.03398   0.02552  -0.0899   0.0989   0.0453
   4.000   0.9709   0.03572   0.02735  -0.0895   0.0859   0.0466
   4.250   0.9935   0.03778   0.02960  -0.0889   0.0763   0.0546
   4.500   1.0132   0.03834   0.03153  -0.0878   0.0720   1.0000
   4.750   1.0345   0.04051   0.03392  -0.0867   0.0652   1.0000
   5.000   1.0565   0.04269   0.03608  -0.0861   0.0616   1.0000
   5.500   1.0957   0.04764   0.04146  -0.0840   0.0541   1.0000
   5.750   1.1146   0.05048   0.04429  -0.0834   0.0522   1.0000
   6.000   1.1331   0.05528   0.04887  -0.0836   0.0511   1.0000
   6.250   1.1450   0.05648   0.05088  -0.0808   0.0472   1.0000
   6.500   1.1596   0.05937   0.05389  -0.0798   0.0454   1.0000
   6.750   1.1739   0.06288   0.05735  -0.0794   0.0444   1.0000
   7.000   1.1916   0.06945   0.06360  -0.0802   0.0436   1.0000
   7.250   1.1843   0.07031   0.06554  -0.0761   0.0407   1.0000
   7.500   1.1895   0.07386   0.06924  -0.0751   0.0395   1.0000
   7.750   1.1935   0.07749   0.07295  -0.0741   0.0388   1.0000
   8.000   1.1975   0.08128   0.07678  -0.0733   0.0383   1.0000
   8.250   1.2067   0.08567   0.08107  -0.0729   0.0379   1.0000
   8.500   1.2127   0.09145   0.08680  -0.0728   0.0375   1.0000
   8.750   1.1904   0.09523   0.09093  -0.0711   0.0374   1.0000
   9.000   1.1678   0.09913   0.09501  -0.0697   0.0374   1.0000
   9.250   1.1466   0.10366   0.09967  -0.0701   0.0374   1.0000
<< Back to Eiffel 10 (Wright) - 1903 Wright Flyer airfoil (eiffel10-il)

Polar data table (+)

Polar graphs


<< Back to Eiffel 10 (Wright) - 1903 Wright Flyer airfoil (eiffel10-il)