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=100,000 Ncrit=9


Details Polar file
Airfoil: Eiffel 10 (Wright) - 1903 Wright Flyer airfoil (eiffel10-il)
Reynolds number: 100,000
Max Cl/Cd: 27.89 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-eiffel10-il-100000.txt
Download as CSV file: xf-eiffel10-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.3392   0.11604   0.11077  -0.0094   1.0000   0.0359
  -9.750  -0.3354   0.11379   0.10855  -0.0107   1.0000   0.0361
  -9.500  -0.3309   0.11117   0.10598  -0.0119   1.0000   0.0362
  -9.250  -0.3210   0.10659   0.10142  -0.0115   1.0000   0.0367
  -9.000  -0.3142   0.10332   0.09819  -0.0120   1.0000   0.0373
  -8.750  -0.3078   0.10030   0.09520  -0.0126   1.0000   0.0379
  -8.500  -0.3017   0.09736   0.09231  -0.0134   1.0000   0.0386
  -8.250  -0.2958   0.09451   0.08950  -0.0142   1.0000   0.0392
  -8.000  -0.2902   0.09176   0.08680  -0.0151   1.0000   0.0399
  -7.750  -0.2853   0.08928   0.08437  -0.0161   1.0000   0.0405
  -7.500  -0.2826   0.08736   0.08251  -0.0175   1.0000   0.0410
  -7.250  -0.2794   0.08579   0.08102  -0.0201   1.0000   0.0412
  -7.000  -0.2710   0.08114   0.07643  -0.0187   1.0000   0.0419
  -6.750  -0.2637   0.07783   0.07319  -0.0187   1.0000   0.0429
  -6.500  -0.2570   0.07506   0.07049  -0.0194   1.0000   0.0440
  -6.250  -0.2513   0.07258   0.06809  -0.0202   1.0000   0.0452
  -6.000  -0.2470   0.07058   0.06617  -0.0213   1.0000   0.0462
  -5.750  -0.2437   0.06928   0.06495  -0.0228   1.0000   0.0468
  -5.500  -0.2380   0.06865   0.06435  -0.0261   1.0000   0.0471
  -5.250  -0.2376   0.06608   0.06186  -0.0246   1.0000   0.0474
  -5.000  -0.2400   0.06316   0.05903  -0.0213   1.0000   0.0480
  -4.750  -0.2374   0.06092   0.05685  -0.0204   1.0000   0.0489
  -4.500  -0.2307   0.05883   0.05481  -0.0208   1.0000   0.0500
  -4.250  -0.2197   0.05677   0.05276  -0.0224   1.0000   0.0515
  -4.000  -0.1969   0.05515   0.05112  -0.0274   1.0000   0.0535
  -3.750  -0.1261   0.06163   0.05707  -0.0469   1.0000   0.0549
  -3.500  -0.1185   0.05912   0.05466  -0.0456   1.0000   0.0562
  -3.250  -0.1004   0.05700   0.05258  -0.0469   1.0000   0.0585
  -3.000  -0.0009   0.05155   0.04702  -0.0652   0.9863   0.0645
  -2.750   0.1052   0.04634   0.04168  -0.0832   0.9650   0.0739
  -2.250   0.1993   0.03960   0.03476  -0.0915   0.8652   0.0878
  -2.000   0.2388   0.03921   0.03382  -0.0931   0.7882   0.0983
  -1.750   0.2493   0.03644   0.03065  -0.0897   0.6551   0.1019
  -1.500   0.2780   0.03851   0.03001  -0.0913   0.0815   0.1132
  -1.250   0.3034   0.03639   0.02798  -0.0917   0.0736   0.1227
  -1.000   0.3362   0.03488   0.02645  -0.0933   0.0706   0.1369
  -0.750   0.3726   0.03395   0.02540  -0.0954   0.0692   0.1602
  -0.500   0.3995   0.03232   0.02392  -0.0958   0.0686   0.1818
   0.500   0.5008   0.02808   0.01996  -0.0947   0.0698   0.4255
   0.750   0.5298   0.02691   0.01881  -0.0943   0.0710   0.4769
   1.000   0.5659   0.02667   0.01840  -0.0952   0.0731   0.4848
   1.250   0.6242   0.02836   0.01910  -0.0986   0.0755   0.2660
   1.500   0.6674   0.02836   0.01823  -0.0976   0.0773   0.1199
   1.750   0.6997   0.02886   0.01828  -0.0965   0.0800   0.0953
   2.000   0.7308   0.02910   0.01826  -0.0955   0.0831   0.0839
   2.250   0.7620   0.02898   0.01833  -0.0944   0.0888   0.0790
   2.500   0.7923   0.03007   0.01930  -0.0935   0.0927   0.0743
   2.750   0.8232   0.03046   0.01980  -0.0923   0.0991   0.0731
   3.000   0.8541   0.03180   0.02122  -0.0911   0.1118   0.0760
   3.250   0.8891   0.03291   0.02283  -0.0893   0.1399   0.0794
   3.500   0.9314   0.03340   0.02401  -0.0874   0.1920   0.0823
   4.500   1.0025   0.07158   0.06715  -0.1317   0.4996   0.0816
   4.750   1.0890   0.05732   0.05334  -0.1018   0.4029   1.0000
   5.000   1.1036   0.05475   0.05018  -0.0922   0.3322   1.0000
   5.250   1.1239   0.05502   0.04979  -0.0866   0.2894   1.0000
   6.000   1.1543   0.06226   0.05673  -0.0805   0.1971   1.0000
   6.250   1.1666   0.06525   0.05959  -0.0790   0.1779   1.0000
   6.500   1.1753   0.06863   0.06297  -0.0780   0.1615   1.0000
   6.750   1.1777   0.07248   0.06697  -0.0772   0.1477   1.0000
   7.000   1.2014   0.07954   0.07344  -0.0769   0.1412   1.0000
   7.250   1.1861   0.08032   0.07495  -0.0757   0.1301   1.0000
   7.500   1.2057   0.08720   0.08147  -0.0753   0.1251   1.0000
   7.750   1.1681   0.09023   0.08517  -0.0757   0.1200   1.0000
   8.000   1.1577   0.09481   0.08986  -0.0761   0.1163   1.0000
   8.250   1.2082   0.10109   0.09559  -0.0739   0.1097   1.0000
   8.500   1.1740   0.10473   0.09961  -0.0743   0.1091   1.0000
   8.750   1.1427   0.10962   0.10467  -0.0756   0.1088   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)