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

AG18 (ag18-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: AG18 (ag18-il)
Reynolds number: 500,000
Max Cl/Cd: 81.83 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag18-il-500000.txt
Download as CSV file: xf-ag18-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG18                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -4.250  -0.2410   0.01709   0.01238  -0.0513   1.0000   0.0142
  -4.000  -0.2144   0.01628   0.01135  -0.0509   1.0000   0.0159
  -3.750  -0.1874   0.01403   0.00876  -0.0506   1.0000   0.0156
  -3.500  -0.1604   0.01211   0.00658  -0.0501   1.0000   0.0150
  -3.250  -0.1335   0.01081   0.00512  -0.0497   1.0000   0.0151
  -3.000  -0.1068   0.00991   0.00411  -0.0492   1.0000   0.0159
  -2.750  -0.0792   0.00890   0.00296  -0.0490   1.0000   0.0187
  -2.500  -0.0525   0.00846   0.00248  -0.0487   1.0000   0.0257
  -2.250  -0.0249   0.00784   0.00197  -0.0486   1.0000   0.0603
  -2.000   0.0047   0.00739   0.00179  -0.0492   0.9989   0.1300
  -1.750   0.0445   0.00687   0.00164  -0.0521   0.9937   0.2393
  -1.500   0.0826   0.00642   0.00156  -0.0545   0.9860   0.3545
  -1.250   0.1193   0.00605   0.00147  -0.0564   0.9755   0.4547
  -1.000   0.1540   0.00566   0.00142  -0.0579   0.9614   0.5657
  -0.750   0.1840   0.00511   0.00137  -0.0581   0.9426   0.7249
  -0.500   0.2101   0.00449   0.00126  -0.0567   0.9189   1.0000
  -0.250   0.2362   0.00455   0.00119  -0.0559   0.8915   1.0000
   0.000   0.2619   0.00464   0.00113  -0.0550   0.8624   1.0000
   0.250   0.2877   0.00476   0.00109  -0.0542   0.8322   1.0000
   0.500   0.3140   0.00489   0.00106  -0.0536   0.8011   1.0000
   0.750   0.3405   0.00504   0.00105  -0.0530   0.7692   1.0000
   1.000   0.3671   0.00521   0.00106  -0.0525   0.7365   1.0000
   1.250   0.3939   0.00538   0.00107  -0.0521   0.7022   1.0000
   1.500   0.4207   0.00557   0.00110  -0.0517   0.6662   1.0000
   1.750   0.4476   0.00578   0.00116  -0.0514   0.6276   1.0000
   2.000   0.4744   0.00601   0.00122  -0.0510   0.5862   1.0000
   2.250   0.5013   0.00627   0.00130  -0.0508   0.5429   1.0000
   2.500   0.5281   0.00655   0.00139  -0.0505   0.4982   1.0000
   2.750   0.5550   0.00684   0.00151  -0.0503   0.4539   1.0000
   3.000   0.5818   0.00715   0.00168  -0.0501   0.4116   1.0000
   3.250   0.6087   0.00746   0.00183  -0.0500   0.3719   1.0000
   3.500   0.6356   0.00778   0.00201  -0.0498   0.3357   1.0000
   3.750   0.6624   0.00810   0.00220  -0.0497   0.3015   1.0000
   4.000   0.6892   0.00843   0.00243  -0.0495   0.2692   1.0000
   4.250   0.7160   0.00875   0.00266  -0.0494   0.2404   1.0000
   4.500   0.7427   0.00910   0.00290  -0.0492   0.2101   1.0000
   4.750   0.7694   0.00946   0.00318  -0.0490   0.1837   1.0000
   5.000   0.7958   0.00985   0.00347  -0.0489   0.1569   1.0000
   5.250   0.8222   0.01026   0.00383  -0.0487   0.1328   1.0000
   5.500   0.8484   0.01071   0.00419  -0.0485   0.1103   1.0000
   5.750   0.8746   0.01114   0.00457  -0.0483   0.0910   1.0000
   6.000   0.9005   0.01163   0.00501  -0.0480   0.0739   1.0000
   6.250   0.9262   0.01219   0.00551  -0.0477   0.0582   1.0000
   6.500   0.9518   0.01272   0.00604  -0.0474   0.0456   1.0000
   6.750   0.9771   0.01335   0.00670  -0.0471   0.0342   1.0000
   7.000   1.0017   0.01416   0.00751  -0.0466   0.0243   1.0000
   7.250   1.0270   0.01474   0.00812  -0.0462   0.0182   1.0000
   7.500   1.0503   0.01591   0.00945  -0.0455   0.0138   1.0000
   7.750   1.0741   0.01685   0.01048  -0.0449   0.0112   1.0000
   8.000   1.0914   0.01944   0.01334  -0.0435   0.0091   1.0000
   8.250   1.1163   0.01991   0.01390  -0.0431   0.0083   1.0000
   8.500   1.1382   0.02117   0.01533  -0.0423   0.0075   1.0000
   8.750   1.1589   0.02264   0.01702  -0.0414   0.0069   1.0000
   9.000   1.1790   0.02410   0.01866  -0.0406   0.0064   1.0000
   9.250   1.1972   0.02586   0.02059  -0.0396   0.0059   1.0000
   9.500   1.2083   0.02917   0.02425  -0.0381   0.0055   1.0000
   9.750   1.2109   0.03401   0.02960  -0.0360   0.0053   1.0000
  10.000   1.2086   0.03898   0.03506  -0.0341   0.0052   1.0000
  10.250   1.2043   0.04334   0.03980  -0.0325   0.0052   1.0000
  10.500   1.1967   0.04721   0.04396  -0.0310   0.0052   1.0000
  10.750   1.1838   0.05074   0.04770  -0.0297   0.0053   1.0000
  11.000   1.1699   0.05515   0.05231  -0.0310   0.0053   1.0000
  11.250   1.1568   0.06083   0.05819  -0.0350   0.0053   1.0000
  11.500   1.1453   0.06772   0.06524  -0.0409   0.0053   1.0000
  11.750   1.1331   0.07592   0.07360  -0.0479   0.0053   1.0000
  12.000   1.1209   0.08455   0.08235  -0.0545   0.0054   1.0000
  12.250   1.1069   0.09356   0.09146  -0.0604   0.0054   1.0000
  12.500   1.0933   0.10229   0.10027  -0.0655   0.0055   1.0000
  12.750   1.0750   0.11226   0.11032  -0.0707   0.0056   1.0000
<< Back to AG18 (ag18-il)

Polar data table (+)

Polar graphs


<< Back to AG18 (ag18-il)