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E182 (8.47%) (e182-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: E182 (8.47%) (e182-il)
Reynolds number: 200,000
Max Cl/Cd: 64.1 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e182-il-200000.txt
Download as CSV file: xf-e182-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E182 (8.47%)                                    
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.4820   0.09802   0.09512   0.0131   1.0000   0.0473
  -9.000  -0.4810   0.09371   0.09082   0.0114   1.0000   0.0487
  -8.750  -0.5882   0.09467   0.09161   0.0089   1.0000   0.0443
  -8.500  -0.5769   0.09217   0.08909   0.0101   1.0000   0.0464
  -8.250  -0.5748   0.08837   0.08532   0.0079   1.0000   0.0476
  -8.000  -0.5756   0.08423   0.08123   0.0049   1.0000   0.0488
  -7.750  -0.5795   0.07975   0.07680   0.0008   1.0000   0.0499
  -7.500  -0.5794   0.07422   0.07126  -0.0051   1.0000   0.0516
  -7.250  -0.5817   0.06745   0.06413  -0.0152   1.0000   0.0545
  -6.750  -0.5656   0.05638   0.05299  -0.0179   1.0000   0.0568
  -6.500  -0.5495   0.05337   0.04999  -0.0184   1.0000   0.0587
  -6.250  -0.5327   0.04986   0.04637  -0.0197   1.0000   0.0618
  -6.000  -0.5188   0.04461   0.04049  -0.0221   1.0000   0.0691
  -5.750  -0.4969   0.04128   0.03724  -0.0231   1.0000   0.0712
  -5.500  -0.4692   0.03826   0.03404  -0.0252   0.9776   0.0760
  -5.250  -0.4381   0.02750   0.02188  -0.0250   0.9526   0.0383
  -5.000  -0.4110   0.02455   0.01805  -0.0232   0.9322   0.0315
  -4.750  -0.3910   0.02166   0.01476  -0.0215   0.9148   0.0306
  -4.500  -0.3695   0.01987   0.01263  -0.0197   0.8995   0.0305
  -4.250  -0.3467   0.01897   0.01145  -0.0182   0.8862   0.0320
  -4.000  -0.3245   0.01705   0.00936  -0.0168   0.8747   0.0335
  -3.750  -0.3008   0.01588   0.00811  -0.0156   0.8632   0.0343
  -3.500  -0.2770   0.01498   0.00713  -0.0145   0.8526   0.0362
  -3.250  -0.2534   0.01427   0.00633  -0.0134   0.8433   0.0393
  -3.000  -0.2294   0.01350   0.00547  -0.0123   0.8339   0.0461
  -2.750  -0.2076   0.01207   0.00464  -0.0112   0.8248   0.1652
  -2.500  -0.1867   0.01118   0.00437  -0.0102   0.8171   0.3248
  -2.250  -0.1672   0.01024   0.00420  -0.0088   0.8080   0.5196
  -2.000  -0.1489   0.00911   0.00427  -0.0055   0.8012   0.8213
  -1.750  -0.0664   0.00922   0.00432  -0.0153   0.7949   0.9642
  -1.500  -0.0008   0.00928   0.00409  -0.0229   0.7891   0.9982
  -1.250   0.0276   0.00927   0.00394  -0.0233   0.7804   1.0000
  -1.000   0.0513   0.00929   0.00381  -0.0225   0.7734   1.0000
  -0.750   0.0762   0.00932   0.00371  -0.0221   0.7651   1.0000
  -0.500   0.1006   0.00938   0.00364  -0.0214   0.7584   1.0000
  -0.250   0.1259   0.00943   0.00361  -0.0210   0.7506   1.0000
   0.000   0.1507   0.00951   0.00358  -0.0204   0.7441   1.0000
   0.250   0.1763   0.00958   0.00359  -0.0200   0.7365   1.0000
   0.500   0.2014   0.00967   0.00360  -0.0194   0.7302   1.0000
   0.750   0.2273   0.00976   0.00365  -0.0191   0.7225   1.0000
   1.000   0.2523   0.00986   0.00368  -0.0185   0.7164   1.0000
   1.250   0.2785   0.00996   0.00377  -0.0182   0.7084   1.0000
   1.500   0.3037   0.01006   0.00381  -0.0176   0.7019   1.0000
   1.750   0.3296   0.01015   0.00390  -0.0172   0.6933   1.0000
   2.000   0.3551   0.01025   0.00397  -0.0166   0.6856   1.0000
   2.250   0.3806   0.01032   0.00404  -0.0160   0.6770   1.0000
   2.500   0.4064   0.01041   0.00414  -0.0156   0.6678   1.0000
   2.750   0.4315   0.01047   0.00416  -0.0147   0.6596   1.0000
   3.000   0.4574   0.01052   0.00425  -0.0143   0.6488   1.0000
   3.250   0.4831   0.01058   0.00436  -0.0137   0.6382   1.0000
   3.500   0.5086   0.01062   0.00441  -0.0130   0.6277   1.0000
   3.750   0.5340   0.01064   0.00443  -0.0123   0.6166   1.0000
   4.000   0.5597   0.01065   0.00449  -0.0116   0.6035   1.0000
   4.250   0.5855   0.01065   0.00458  -0.0110   0.5895   1.0000
   4.500   0.6110   0.01063   0.00462  -0.0103   0.5720   1.0000
   4.750   0.6360   0.01052   0.00452  -0.0094   0.5427   1.0000
   5.000   0.6609   0.01051   0.00448  -0.0086   0.5043   1.0000
   5.250   0.6852   0.01069   0.00456  -0.0078   0.4459   1.0000
   5.500   0.7064   0.01152   0.00482  -0.0071   0.3203   1.0000
   5.750   0.7254   0.01304   0.00564  -0.0067   0.1864   1.0000
   6.000   0.7401   0.01550   0.00717  -0.0059   0.0452   1.0000
   6.250   0.7610   0.01669   0.00842  -0.0049   0.0353   1.0000
   6.500   0.7817   0.01774   0.00952  -0.0041   0.0304   1.0000
   6.750   0.8001   0.01911   0.01095  -0.0029   0.0278   1.0000
   7.000   0.8197   0.02031   0.01224  -0.0018   0.0265   1.0000
   7.250   0.8393   0.02165   0.01366  -0.0005   0.0257   1.0000
   7.500   0.8594   0.02318   0.01527   0.0007   0.0250   1.0000
   7.750   0.8805   0.02491   0.01715   0.0019   0.0248   1.0000
   8.000   0.9020   0.02703   0.01945   0.0031   0.0249   1.0000
   8.250   0.9229   0.02954   0.02224   0.0042   0.0254   1.0000
   8.500   0.9415   0.03256   0.02560   0.0054   0.0263   1.0000
   8.750   0.9563   0.03639   0.02981   0.0067   0.0274   1.0000
   9.000   0.9680   0.04211   0.03579   0.0077   0.0288   1.0000
   9.250   0.9800   0.04484   0.03883   0.0089   0.0290   1.0000
   9.500   0.9932   0.04626   0.04060   0.0105   0.0299   1.0000
  13.500   0.7857   0.15037   0.14672  -0.0257   0.0496   1.0000
  13.750   0.7837   0.15604   0.15237  -0.0295   0.0494   1.0000
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