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

E71 (5.15%) (e71-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: E71 (5.15%) (e71-il)
Reynolds number: 200,000
Max Cl/Cd: 108.42 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e71-il-200000.txt
Download as CSV file: xf-e71-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E71  (5.15%)                                    
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.3796   0.10200   0.09866  -0.0201   1.0000   0.0245
  -7.250  -0.3888   0.10111   0.09784  -0.0189   1.0000   0.0248
  -7.000  -0.3978   0.10004   0.09685  -0.0176   1.0000   0.0250
  -6.750  -0.3982   0.09794   0.09479  -0.0179   1.0000   0.0251
  -6.500  -0.3961   0.09567   0.09256  -0.0196   1.0000   0.0253
  -6.250  -0.3913   0.09296   0.08989  -0.0217   1.0000   0.0254
  -6.000  -0.3858   0.08954   0.08653  -0.0238   1.0000   0.0255
  -5.750  -0.3963   0.08481   0.08183  -0.0165   1.0000   0.0273
  -5.500  -0.3928   0.08251   0.07955  -0.0155   1.0000   0.0292
  -5.250  -0.3863   0.07979   0.07685  -0.0169   1.0000   0.0313
  -5.000  -0.3754   0.07662   0.07370  -0.0201   1.0000   0.0332
  -4.750  -0.3381   0.07226   0.06928  -0.0334   1.0000   0.0364
  -4.500  -0.2943   0.06747   0.06432  -0.0459   1.0000   0.0371
  -4.250  -0.2772   0.06066   0.05753  -0.0498   0.9988   0.0384
  -4.000  -0.2549   0.05765   0.05452  -0.0511   0.9957   0.0407
  -3.750  -0.2086   0.05304   0.04979  -0.0604   0.9929   0.0453
  -3.500  -0.1305   0.04462   0.04099  -0.0793   0.9919   0.0521
  -3.250  -0.0976   0.04194   0.03828  -0.0828   0.9890   0.0566
  -3.000  -0.0394   0.03694   0.03295  -0.0930   0.9881   0.0679
  -2.750   0.0134   0.03315   0.02878  -0.1008   0.9872   0.0808
  -2.500   0.0624   0.03020   0.02552  -0.1069   0.9860   0.0941
  -2.250   0.1075   0.02792   0.02302  -0.1117   0.9847   0.1089
  -2.000   0.1781   0.02050   0.01402  -0.1178   0.9885   0.0415
  -1.750   0.2199   0.01950   0.01256  -0.1199   0.9864   0.0375
  -1.500   0.2608   0.01832   0.01112  -0.1225   0.9848   0.0373
  -1.250   0.3020   0.01706   0.00983  -0.1254   0.9836   0.0430
  -1.000   0.3397   0.01645   0.00918  -0.1275   0.9798   0.0446
  -0.750   0.3787   0.01605   0.00870  -0.1298   0.9755   0.0483
  -0.500   0.4223   0.01529   0.00833  -0.1333   0.9731   0.1697
  -0.250   0.4639   0.01418   0.00865  -0.1367   0.9715   0.6391
   0.000   0.4870   0.01354   0.00850  -0.1351   0.9634   1.0000
   0.250   0.5284   0.01361   0.00843  -0.1379   0.9590   1.0000
   0.500   0.5628   0.01361   0.00834  -0.1393   0.9515   1.0000
   0.750   0.6047   0.01354   0.00819  -0.1420   0.9466   1.0000
   1.000   0.6397   0.01344   0.00805  -0.1434   0.9385   1.0000
   1.250   0.6828   0.01321   0.00780  -0.1462   0.9335   1.0000
   1.500   0.7177   0.01302   0.00764  -0.1474   0.9248   1.0000
   1.750   0.7631   0.01259   0.00724  -0.1504   0.9199   1.0000
   2.000   0.7985   0.01226   0.00694  -0.1515   0.9102   1.0000
   2.250   0.8456   0.01163   0.00639  -0.1547   0.9059   1.0000
   2.500   0.8815   0.01123   0.00611  -0.1557   0.8949   1.0000
   2.750   0.9196   0.01077   0.00574  -0.1570   0.8826   1.0000
   3.000   0.9588   0.01032   0.00539  -0.1585   0.8675   1.0000
   3.250   0.9945   0.01002   0.00517  -0.1593   0.8450   1.0000
   3.500   1.0306   0.00981   0.00504  -0.1602   0.8147   1.0000
   3.750   1.0625   0.00980   0.00499  -0.1602   0.7723   1.0000
   4.000   1.0893   0.01006   0.00504  -0.1591   0.7038   1.0000
   4.250   1.1052   0.01097   0.00524  -0.1558   0.5715   1.0000
   4.500   1.1134   0.01252   0.00584  -0.1519   0.4137   1.0000
   4.750   1.1285   0.01376   0.00653  -0.1496   0.3076   1.0000
   5.000   1.1433   0.01530   0.00742  -0.1476   0.1877   1.0000
   5.250   1.1541   0.01787   0.00891  -0.1450   0.0398   1.0000
   5.500   1.1735   0.01936   0.01041  -0.1431   0.0237   1.0000
   5.750   1.1937   0.02060   0.01179  -0.1414   0.0203   1.0000
   6.000   1.2110   0.02239   0.01370  -0.1393   0.0186   1.0000
   6.250   1.2285   0.02480   0.01624  -0.1372   0.0179   1.0000
   6.500   1.2508   0.02761   0.01919  -0.1359   0.0177   1.0000
   6.750   1.2759   0.03020   0.02197  -0.1350   0.0178   1.0000
   7.000   1.3006   0.03243   0.02449  -0.1339   0.0182   1.0000
   7.250   1.3246   0.03450   0.02688  -0.1325   0.0189   1.0000
   7.500   1.3468   0.03949   0.03270  -0.1298   0.0219   1.0000
   7.750   1.3588   0.04577   0.03962  -0.1266   0.0252   1.0000
  15.750   0.9677   0.22979   0.22760  -0.1588   0.0301   1.0000
  16.000   0.9713   0.23374   0.23156  -0.1602   0.0295   1.0000
<< Back to E71 (5.15%) (e71-il)

Polar data table (+)

Polar graphs


<< Back to E71 (5.15%) (e71-il)