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

AG19 (ag19-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: AG19 (ag19-il)
Reynolds number: 200,000
Max Cl/Cd: 60.39 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag19-il-200000-n5.txt
Download as CSV file: xf-ag19-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG19                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.4199   0.07386   0.07068  -0.0026   1.0000   0.0153
  -7.000  -0.4193   0.06970   0.06656  -0.0039   1.0000   0.0141
  -6.750  -0.4178   0.06516   0.06205  -0.0064   1.0000   0.0133
  -6.500  -0.4130   0.05965   0.05656  -0.0109   1.0000   0.0124
  -6.250  -0.4051   0.05292   0.04983  -0.0174   1.0000   0.0115
  -6.000  -0.3907   0.04374   0.04058  -0.0269   1.0000   0.0105
  -5.750  -0.3728   0.03688   0.03361  -0.0332   1.0000   0.0104
  -5.500  -0.3539   0.03095   0.02752  -0.0382   1.0000   0.0108
  -5.250  -0.3519   0.04120   0.03728  -0.0432   1.0000   0.0117
  -5.000  -0.3253   0.03741   0.03328  -0.0459   1.0000   0.0135
  -4.750  -0.2957   0.03156   0.02699  -0.0488   1.0000   0.0135
  -4.500  -0.2660   0.02608   0.02092  -0.0507   1.0000   0.0129
  -4.250  -0.2368   0.02189   0.01604  -0.0516   1.0000   0.0126
  -4.000  -0.2084   0.01901   0.01260  -0.0517   1.0000   0.0126
  -3.750  -0.1807   0.01695   0.01013  -0.0515   1.0000   0.0128
  -3.500  -0.1535   0.01537   0.00826  -0.0512   1.0000   0.0133
  -3.250  -0.1266   0.01410   0.00679  -0.0508   1.0000   0.0143
  -3.000  -0.0999   0.01308   0.00563  -0.0504   1.0000   0.0157
  -2.750  -0.0735   0.01225   0.00474  -0.0501   1.0000   0.0197
  -2.500  -0.0469   0.01164   0.00404  -0.0498   1.0000   0.0242
  -2.250  -0.0199   0.01102   0.00337  -0.0495   1.0000   0.0329
  -2.000   0.0068   0.01054   0.00293  -0.0493   1.0000   0.0545
  -1.500   0.0706   0.00955   0.00249  -0.0516   0.9912   0.1991
  -1.250   0.1080   0.00903   0.00234  -0.0539   0.9779   0.3191
  -1.000   0.1435   0.00849   0.00226  -0.0556   0.9620   0.4649
  -0.750   0.1760   0.00781   0.00219  -0.0564   0.9434   0.6511
  -0.500   0.2082   0.00705   0.00202  -0.0563   0.9215   1.0000
  -0.250   0.2402   0.00708   0.00188  -0.0567   0.8955   1.0000
   0.000   0.2693   0.00713   0.00179  -0.0565   0.8667   1.0000
   0.250   0.2964   0.00722   0.00171  -0.0559   0.8360   1.0000
   0.500   0.3225   0.00734   0.00167  -0.0550   0.8043   1.0000
   0.750   0.3484   0.00748   0.00164  -0.0542   0.7711   1.0000
   1.000   0.3743   0.00765   0.00163  -0.0534   0.7368   1.0000
   1.250   0.4003   0.00784   0.00164  -0.0527   0.7005   1.0000
   1.500   0.4263   0.00806   0.00167  -0.0521   0.6616   1.0000
   1.750   0.4523   0.00830   0.00174  -0.0514   0.6193   1.0000
   2.000   0.4782   0.00858   0.00181  -0.0509   0.5740   1.0000
   2.250   0.5042   0.00890   0.00190  -0.0504   0.5262   1.0000
   2.500   0.5302   0.00924   0.00203  -0.0499   0.4788   1.0000
   2.750   0.5563   0.00959   0.00218  -0.0496   0.4338   1.0000
   3.000   0.5825   0.00996   0.00241  -0.0492   0.3913   1.0000
   3.250   0.6088   0.01034   0.00263  -0.0490   0.3529   1.0000
   3.500   0.6351   0.01070   0.00286  -0.0487   0.3176   1.0000
   3.750   0.6615   0.01109   0.00313  -0.0485   0.2858   1.0000
   4.000   0.6878   0.01146   0.00341  -0.0483   0.2559   1.0000
   4.250   0.7142   0.01185   0.00377  -0.0481   0.2282   1.0000
   4.500   0.7404   0.01226   0.00411  -0.0479   0.2020   1.0000
   4.750   0.7665   0.01270   0.00448  -0.0476   0.1765   1.0000
   5.000   0.7926   0.01313   0.00489  -0.0474   0.1540   1.0000
   5.250   0.8185   0.01362   0.00533  -0.0471   0.1322   1.0000
   5.500   0.8442   0.01413   0.00587  -0.0469   0.1138   1.0000
   5.750   0.8698   0.01468   0.00640  -0.0466   0.0970   1.0000
   6.000   0.8953   0.01523   0.00698  -0.0463   0.0817   1.0000
   6.250   0.9205   0.01584   0.00763  -0.0459   0.0693   1.0000
   6.500   0.9455   0.01649   0.00831  -0.0456   0.0567   1.0000
   6.750   0.9701   0.01724   0.00909  -0.0452   0.0464   1.0000
   7.000   0.9943   0.01807   0.00999  -0.0448   0.0361   1.0000
   7.250   1.0187   0.01886   0.01093  -0.0443   0.0289   1.0000
   7.500   1.0420   0.01986   0.01203  -0.0438   0.0223   1.0000
   7.750   1.0645   0.02106   0.01335  -0.0431   0.0178   1.0000
   8.000   1.0858   0.02251   0.01501  -0.0423   0.0148   1.0000
   8.250   1.1083   0.02353   0.01619  -0.0417   0.0119   1.0000
   8.500   1.1276   0.02528   0.01812  -0.0409   0.0103   1.0000
   8.750   1.1438   0.02776   0.02088  -0.0396   0.0094   1.0000
   9.000   1.1612   0.02992   0.02336  -0.0385   0.0088   1.0000
   9.250   1.1763   0.03252   0.02639  -0.0373   0.0083   1.0000
   9.500   1.1892   0.03530   0.02956  -0.0360   0.0078   1.0000
   9.750   1.2018   0.03763   0.03221  -0.0351   0.0071   1.0000
  10.000   1.2145   0.03943   0.03423  -0.0343   0.0064   1.0000
  10.250   1.2229   0.04164   0.03664  -0.0335   0.0059   1.0000
  10.500   1.2197   0.04543   0.04076  -0.0323   0.0055   1.0000
  10.750   1.2074   0.04964   0.04530  -0.0309   0.0054   1.0000
  11.000   1.1909   0.05420   0.05013  -0.0310   0.0054   1.0000
  11.250   1.1754   0.05975   0.05593  -0.0340   0.0054   1.0000
  11.500   1.1615   0.06660   0.06300  -0.0394   0.0054   1.0000
  11.750   1.1483   0.07473   0.07132  -0.0463   0.0055   1.0000
  12.000   1.1329   0.08421   0.08099  -0.0535   0.0056   1.0000
  12.250   1.0397   0.11517   0.11219  -0.0709   0.0083   1.0000
  12.500   1.0148   0.12682   0.12380  -0.0764   0.0086   1.0000
<< Back to AG19 (ag19-il)

Polar data table (+)

Polar graphs


<< Back to AG19 (ag19-il)