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

AG13 (ag13-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: AG13 (ag13-il)
Reynolds number: 500,000
Max Cl/Cd: 79.92 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag13-il-500000.txt
Download as CSV file: xf-ag13-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG13                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5591   0.09934   0.09715   0.0180   1.0000   0.0125
  -8.000  -0.5546   0.09560   0.09343   0.0160   1.0000   0.0128
  -7.750  -0.5501   0.09180   0.08965   0.0137   1.0000   0.0131
  -7.500  -0.5454   0.08793   0.08580   0.0111   1.0000   0.0135
  -7.250  -0.5386   0.08374   0.08163   0.0073   1.0000   0.0139
  -7.000  -0.5266   0.07886   0.07677   0.0015   1.0000   0.0145
  -6.750  -0.5088   0.07337   0.07127  -0.0061   1.0000   0.0153
  -6.500  -0.4842   0.06815   0.06603  -0.0139   1.0000   0.0160
  -6.250  -0.4605   0.06211   0.05994  -0.0216   1.0000   0.0162
  -6.000  -0.4356   0.05542   0.05314  -0.0288   1.0000   0.0162
  -4.500  -0.2864   0.01856   0.01399  -0.0476   1.0000   0.0140
  -4.250  -0.2589   0.01631   0.01137  -0.0475   1.0000   0.0144
  -4.000  -0.2318   0.01548   0.01033  -0.0471   1.0000   0.0157
  -3.750  -0.2044   0.01304   0.00756  -0.0468   1.0000   0.0151
  -3.500  -0.1773   0.01150   0.00582  -0.0464   1.0000   0.0150
  -3.250  -0.1503   0.01043   0.00461  -0.0459   1.0000   0.0155
  -3.000  -0.1232   0.00955   0.00363  -0.0455   1.0000   0.0168
  -2.750  -0.0958   0.00865   0.00265  -0.0453   1.0000   0.0223
  -2.500  -0.0686   0.00818   0.00215  -0.0450   1.0000   0.0368
  -2.250  -0.0414   0.00762   0.00180  -0.0449   1.0000   0.0888
  -2.000  -0.0145   0.00720   0.00166  -0.0449   1.0000   0.1598
  -1.750   0.0125   0.00681   0.00157  -0.0449   1.0000   0.2467
  -1.500   0.0396   0.00647   0.00152  -0.0450   1.0000   0.3382
  -1.250   0.0792   0.00607   0.00145  -0.0477   0.9935   0.4486
  -1.000   0.1177   0.00567   0.00141  -0.0501   0.9832   0.5628
  -0.750   0.1529   0.00511   0.00137  -0.0516   0.9687   0.7193
  -0.500   0.1843   0.00444   0.00127  -0.0514   0.9463   1.0000
  -0.250   0.2142   0.00446   0.00117  -0.0514   0.9145   1.0000
   0.000   0.2398   0.00454   0.00110  -0.0504   0.8754   1.0000
   0.250   0.2648   0.00469   0.00105  -0.0494   0.8337   1.0000
   0.500   0.2904   0.00487   0.00102  -0.0485   0.7918   1.0000
   0.750   0.3165   0.00506   0.00102  -0.0479   0.7500   1.0000
   1.000   0.3429   0.00527   0.00102  -0.0474   0.7085   1.0000
   1.250   0.3696   0.00549   0.00105  -0.0469   0.6677   1.0000
   1.500   0.3965   0.00570   0.00109  -0.0466   0.6280   1.0000
   1.750   0.4235   0.00592   0.00115  -0.0463   0.5896   1.0000
   2.000   0.4506   0.00615   0.00122  -0.0460   0.5519   1.0000
   2.250   0.4776   0.00639   0.00130  -0.0458   0.5135   1.0000
   2.500   0.5047   0.00663   0.00140  -0.0456   0.4763   1.0000
   2.750   0.5319   0.00688   0.00151  -0.0454   0.4398   1.0000
   3.000   0.5589   0.00715   0.00166  -0.0453   0.4033   1.0000
   3.250   0.5859   0.00744   0.00180  -0.0451   0.3665   1.0000
   3.500   0.6130   0.00772   0.00197  -0.0450   0.3320   1.0000
   3.750   0.6399   0.00804   0.00215  -0.0448   0.2971   1.0000
   4.000   0.6668   0.00836   0.00238  -0.0447   0.2644   1.0000
   4.250   0.6937   0.00868   0.00261  -0.0445   0.2342   1.0000
   4.500   0.7203   0.00906   0.00286  -0.0444   0.2029   1.0000
   4.750   0.7470   0.00942   0.00314  -0.0442   0.1756   1.0000
   5.000   0.7735   0.00983   0.00344  -0.0440   0.1478   1.0000
   5.250   0.7999   0.01024   0.00379  -0.0438   0.1236   1.0000
   5.500   0.8263   0.01067   0.00416  -0.0436   0.1013   1.0000
   5.750   0.8523   0.01116   0.00458  -0.0434   0.0820   1.0000
   6.000   0.8785   0.01161   0.00501  -0.0432   0.0660   1.0000
   6.250   0.9044   0.01212   0.00548  -0.0429   0.0513   1.0000
   6.500   0.9301   0.01269   0.00601  -0.0426   0.0385   1.0000
   6.750   0.9555   0.01334   0.00671  -0.0423   0.0283   1.0000
   7.000   0.9804   0.01410   0.00751  -0.0418   0.0201   1.0000
   7.250   1.0036   0.01538   0.00889  -0.0411   0.0144   1.0000
   7.500   1.0283   0.01619   0.00982  -0.0406   0.0122   1.0000
   7.750   1.0518   0.01725   0.01098  -0.0400   0.0104   1.0000
   8.000   1.0692   0.01996   0.01398  -0.0386   0.0091   1.0000
   8.250   1.0886   0.02214   0.01643  -0.0374   0.0086   1.0000
   8.500   1.1118   0.02305   0.01749  -0.0368   0.0081   1.0000
   8.750   1.1334   0.02432   0.01898  -0.0361   0.0074   1.0000
   9.000   1.1519   0.02638   0.02130  -0.0351   0.0071   1.0000
   9.250   1.1689   0.02863   0.02383  -0.0340   0.0068   1.0000
   9.500   1.1840   0.03104   0.02653  -0.0330   0.0064   1.0000
   9.750   1.1969   0.03355   0.02932  -0.0319   0.0061   1.0000
  10.000   1.2066   0.03634   0.03240  -0.0308   0.0059   1.0000
  10.250   1.2109   0.03967   0.03603  -0.0296   0.0057   1.0000
  10.500   1.2080   0.04359   0.04027  -0.0284   0.0056   1.0000
  10.750   1.1964   0.04770   0.04468  -0.0271   0.0056   1.0000
  11.000   1.1787   0.05224   0.04946  -0.0274   0.0056   1.0000
  11.250   1.1568   0.05948   0.05698  -0.0326   0.0058   1.0000
<< Back to AG13 (ag13-il)

Polar data table (+)

Polar graphs


<< Back to AG13 (ag13-il)