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E182 (8.47%) (e182-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: E182 (8.47%) (e182-il)
Reynolds number: 50,000
Max Cl/Cd: 34.12 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e182-il-50000-n5.txt
Download as CSV file: xf-e182-il-50000-n5.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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5651   0.09276   0.08643   0.0023   1.0000   0.0459
  -8.250  -0.5652   0.08797   0.08172  -0.0011   1.0000   0.0448
  -8.000  -0.5671   0.08297   0.07678  -0.0052   1.0000   0.0437
  -7.750  -0.5675   0.07755   0.07138  -0.0095   1.0000   0.0426
  -7.500  -0.5673   0.07170   0.06547  -0.0136   1.0000   0.0414
  -7.250  -0.5665   0.06553   0.05919  -0.0174   1.0000   0.0401
  -7.000  -0.5646   0.05902   0.05230  -0.0206   1.0000   0.0385
  -6.500  -0.5421   0.05010   0.04248  -0.0224   1.0000   0.0374
  -6.250  -0.5275   0.04613   0.03812  -0.0226   1.0000   0.0374
  -6.000  -0.5116   0.04232   0.03400  -0.0227   1.0000   0.0381
  -5.750  -0.4936   0.03932   0.03076  -0.0227   1.0000   0.0396
  -5.500  -0.4734   0.03678   0.02796  -0.0226   1.0000   0.0417
  -5.250  -0.4519   0.03408   0.02484  -0.0222   1.0000   0.0431
  -5.000  -0.4292   0.03151   0.02185  -0.0217   1.0000   0.0441
  -4.750  -0.4062   0.02926   0.01922  -0.0210   1.0000   0.0455
  -4.500  -0.3842   0.02737   0.01696  -0.0200   1.0000   0.0473
  -4.250  -0.3646   0.02585   0.01515  -0.0185   1.0000   0.0495
  -4.000  -0.3494   0.02460   0.01391  -0.0168   0.9995   0.0534
  -3.750  -0.3113   0.02318   0.01230  -0.0191   0.9830   0.0652
  -3.500  -0.2722   0.02161   0.01066  -0.0214   0.9696   0.0908
  -3.250  -0.2368   0.01971   0.00951  -0.0240   0.9570   0.2096
  -3.000  -0.2081   0.01801   0.00888  -0.0249   0.9441   0.4215
  -2.750  -0.1795   0.01672   0.00875  -0.0234   0.9336   0.6991
  -2.500  -0.0819   0.01636   0.00817  -0.0337   0.9376   0.9523
  -2.250  -0.0182   0.01621   0.00749  -0.0410   0.9308   1.0000
  -2.000   0.0124   0.01627   0.00717  -0.0422   0.9144   1.0000
  -1.750   0.0400   0.01638   0.00697  -0.0426   0.8992   1.0000
  -1.500   0.0656   0.01652   0.00685  -0.0426   0.8854   1.0000
  -1.250   0.0900   0.01669   0.00679  -0.0422   0.8726   1.0000
  -1.000   0.1138   0.01688   0.00675  -0.0416   0.8609   1.0000
  -0.750   0.1369   0.01709   0.00678  -0.0409   0.8491   1.0000
  -0.500   0.1601   0.01731   0.00684  -0.0403   0.8381   1.0000
  -0.250   0.1834   0.01754   0.00694  -0.0396   0.8281   1.0000
   0.000   0.2065   0.01778   0.00705  -0.0387   0.8186   1.0000
   0.250   0.2299   0.01805   0.00723  -0.0381   0.8084   1.0000
   0.500   0.2532   0.01832   0.00742  -0.0374   0.7992   1.0000
   0.750   0.2765   0.01859   0.00763  -0.0366   0.7904   1.0000
   1.000   0.2999   0.01891   0.00793  -0.0360   0.7807   1.0000
   1.250   0.3233   0.01921   0.00819  -0.0352   0.7724   1.0000
   1.500   0.3467   0.01953   0.00852  -0.0345   0.7632   1.0000
   1.750   0.3701   0.01989   0.00889  -0.0339   0.7539   1.0000
   2.000   0.3936   0.02017   0.00920  -0.0329   0.7460   1.0000
   2.250   0.4172   0.02057   0.00966  -0.0324   0.7356   1.0000
   2.500   0.4406   0.02093   0.01008  -0.0316   0.7261   1.0000
   2.750   0.4639   0.02120   0.01044  -0.0304   0.7172   1.0000
   3.000   0.4869   0.02158   0.01093  -0.0297   0.7057   1.0000
   3.250   0.5098   0.02193   0.01139  -0.0287   0.6942   1.0000
   3.500   0.5327   0.02221   0.01178  -0.0275   0.6827   1.0000
   3.750   0.5557   0.02243   0.01216  -0.0261   0.6706   1.0000
   4.000   0.5785   0.02260   0.01248  -0.0246   0.6574   1.0000
   4.250   0.6010   0.02277   0.01280  -0.0231   0.6422   1.0000
   4.500   0.6233   0.02290   0.01312  -0.0216   0.6250   1.0000
   4.750   0.6457   0.02291   0.01332  -0.0198   0.6064   1.0000
   5.000   0.6685   0.02274   0.01336  -0.0177   0.5870   1.0000
   5.250   0.6904   0.02269   0.01354  -0.0159   0.5618   1.0000
   5.500   0.7123   0.02251   0.01358  -0.0139   0.5328   1.0000
   5.750   0.7338   0.02238   0.01366  -0.0119   0.4957   1.0000
   6.000   0.7542   0.02228   0.01372  -0.0096   0.4429   1.0000
   6.250   0.7705   0.02258   0.01368  -0.0069   0.3454   1.0000
   6.500   0.7769   0.02467   0.01462  -0.0049   0.2064   1.0000
   6.750   0.7824   0.02741   0.01644  -0.0038   0.1091   1.0000
   7.000   0.7917   0.02974   0.01846  -0.0024   0.0719   1.0000
   7.250   0.8022   0.03187   0.02047  -0.0009   0.0583   1.0000
   7.500   0.8160   0.03366   0.02246   0.0008   0.0513   1.0000
   7.750   0.8296   0.03556   0.02439   0.0023   0.0469   1.0000
   8.000   0.8492   0.03724   0.02639   0.0038   0.0423   1.0000
   8.250   0.8674   0.03903   0.02833   0.0050   0.0386   1.0000
   8.500   0.8863   0.04128   0.03063   0.0061   0.0363   1.0000
   8.750   0.9077   0.04389   0.03356   0.0072   0.0352   1.0000
   9.000   0.9253   0.04683   0.03691   0.0083   0.0344   1.0000
   9.250   0.9378   0.05006   0.04057   0.0093   0.0339   1.0000
   9.500   0.9448   0.05361   0.04455   0.0104   0.0335   1.0000
   9.750   0.9465   0.05726   0.04857   0.0113   0.0333   1.0000
  10.000   0.9425   0.06105   0.05276   0.0121   0.0333   1.0000
  10.250   0.9322   0.06472   0.05669   0.0129   0.0333   1.0000
  10.500   0.9186   0.06874   0.06093   0.0127   0.0334   1.0000
  10.750   0.9031   0.07329   0.06568   0.0115   0.0336   1.0000
  11.000   0.8857   0.07853   0.07108   0.0091   0.0337   1.0000
  11.250   0.8682   0.08437   0.07699   0.0059   0.0340   1.0000
  11.500   0.8504   0.09099   0.08371   0.0017   0.0343   1.0000
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