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EPPLER 49 AIRFOIL (e49-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 49 AIRFOIL (e49-il)
Reynolds number: 1,000,000
Max Cl/Cd: 155.5 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e49-il-1000000.txt
Download as CSV file: xf-e49-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 49 AIRFOIL                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.0807   0.09201   0.09023  -0.1012   0.9286   0.0069
  -8.250  -0.0728   0.08939   0.08761  -0.1033   0.9281   0.0069
  -8.000  -0.1184   0.09088   0.08916  -0.0895   0.9192   0.0069
  -7.750  -0.1033   0.08710   0.08538  -0.0937   0.9181   0.0069
  -7.500  -0.0837   0.08315   0.08142  -0.0985   0.9174   0.0069
  -7.250  -0.0601   0.07879   0.07705  -0.1043   0.9168   0.0069
  -7.000  -0.0908   0.07683   0.07513  -0.0966   0.9081   0.0071
  -6.750  -0.0746   0.07383   0.07213  -0.0985   0.9072   0.0074
  -6.500  -0.0512   0.07035   0.06863  -0.1033   0.9064   0.0076
  -6.250  -0.0244   0.06658   0.06484  -0.1093   0.9057   0.0079
  -6.000   0.0071   0.06251   0.06073  -0.1163   0.9052   0.0082
  -5.750   0.0043   0.06064   0.05886  -0.1141   0.8976   0.0087
  -5.500   0.0387   0.05642   0.05460  -0.1212   0.8963   0.0091
  -5.250   0.0887   0.05187   0.04996  -0.1313   0.8956   0.0106
  -4.250   0.2912   0.02974   0.02702  -0.1676   0.8953   0.0116
  -4.000   0.3290   0.02758   0.02469  -0.1714   0.8951   0.0121
  -3.750   0.3670   0.02566   0.02257  -0.1748   0.8949   0.0128
  -3.500   0.4052   0.02386   0.02055  -0.1777   0.8947   0.0138
  -3.250   0.4422   0.02250   0.01898  -0.1799   0.8945   0.0156
  -3.000   0.4754   0.02250   0.01888  -0.1806   0.8940   0.0174
  -2.500   0.5531   0.01816   0.01385  -0.1863   0.8940   0.0203
  -2.250   0.5868   0.01749   0.01313  -0.1878   0.8936   0.0226
  -2.000   0.6207   0.01696   0.01252  -0.1892   0.8933   0.0257
  -1.750   0.6542   0.01716   0.01268  -0.1903   0.8928   0.0285
  -1.500   0.6692   0.01640   0.01185  -0.1877   0.8874   0.0281
  -1.250   0.7066   0.01495   0.01027  -0.1894   0.8864   0.0170
  -1.000   0.7459   0.01397   0.00926  -0.1917   0.8850   0.0172
  -0.750   0.7842   0.01311   0.00833  -0.1939   0.8839   0.0196
  -0.500   0.8186   0.01266   0.00785  -0.1954   0.8831   0.0216
  -0.250   0.8527   0.01225   0.00744  -0.1968   0.8824   0.0241
   0.000   0.8881   0.01152   0.00717  -0.1987   0.8818   0.2631
   0.250   0.9201   0.01006   0.00723  -0.2002   0.8814   1.0000
   0.500   0.9556   0.00958   0.00673  -0.2019   0.8807   1.0000
   0.750   0.9714   0.00960   0.00674  -0.1993   0.8735   1.0000
   1.000   1.0022   0.00915   0.00627  -0.2000   0.8696   1.0000
   1.250   1.0371   0.00869   0.00580  -0.2016   0.8671   1.0000
   1.500   1.0479   0.00902   0.00615  -0.1981   0.8567   1.0000
   1.750   1.0725   0.00895   0.00608  -0.1975   0.8483   1.0000
   2.000   1.1066   0.00857   0.00569  -0.1990   0.8394   1.0000
   2.250   1.1443   0.00809   0.00515  -0.2012   0.8187   1.0000
   2.500   1.1880   0.00764   0.00427  -0.2046   0.7411   1.0000
   3.000   1.1938   0.00923   0.00527  -0.1943   0.6354   1.0000
   3.250   1.1990   0.01004   0.00580  -0.1898   0.5767   1.0000
   3.500   1.2027   0.01109   0.00640  -0.1853   0.4910   1.0000
   3.750   1.2106   0.01219   0.00699  -0.1817   0.3948   1.0000
   4.000   1.2218   0.01328   0.00755  -0.1790   0.3002   1.0000
   4.250   1.2345   0.01434   0.00813  -0.1767   0.2123   1.0000
   4.500   1.2421   0.01595   0.00903  -0.1735   0.0885   1.0000
   4.750   1.2529   0.01740   0.01004  -0.1706   0.0077   1.0000
   5.000   1.2736   0.01784   0.01054  -0.1695   0.0066   1.0000
   5.250   1.2936   0.01834   0.01113  -0.1683   0.0063   1.0000
   5.500   1.3127   0.01895   0.01182  -0.1669   0.0059   1.0000
   5.750   1.3324   0.01946   0.01237  -0.1658   0.0056   1.0000
   6.000   1.3510   0.02011   0.01310  -0.1644   0.0054   1.0000
   6.250   1.3689   0.02082   0.01389  -0.1628   0.0051   1.0000
   6.500   1.3858   0.02163   0.01478  -0.1611   0.0050   1.0000
   6.750   1.4017   0.02250   0.01572  -0.1593   0.0046   1.0000
   7.000   1.4155   0.02353   0.01684  -0.1571   0.0043   1.0000
   7.250   1.4273   0.02474   0.01814  -0.1546   0.0041   1.0000
   7.500   1.4375   0.02615   0.01966  -0.1518   0.0041   1.0000
   7.750   1.4463   0.02789   0.02157  -0.1486   0.0043   1.0000
   9.000   1.5164   0.04167   0.03684  -0.1384   0.0060   1.0000
   9.250   1.5100   0.04496   0.04035  -0.1337   0.0060   1.0000
   9.500   1.4997   0.04851   0.04412  -0.1286   0.0059   1.0000
   9.750   1.4856   0.05223   0.04806  -0.1233   0.0059   1.0000
  10.000   1.4691   0.05624   0.05229  -0.1182   0.0059   1.0000
  10.250   1.4489   0.06043   0.05669  -0.1131   0.0059   1.0000
  10.500   1.4234   0.06478   0.06126  -0.1080   0.0059   1.0000
  10.750   1.3825   0.06909   0.06582  -0.1022   0.0059   1.0000
  11.000   1.3543   0.07299   0.06992  -0.0986   0.0059   1.0000
  11.250   1.3275   0.07776   0.07490  -0.0962   0.0059   1.0000
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