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AG08 (ag08-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: AG08 (ag08-il)
Reynolds number: 500,000
Max Cl/Cd: 76.24 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag08-il-500000.txt
Download as CSV file: xf-ag08-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG08                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5792   0.10054   0.09835   0.0184   1.0000   0.0149
  -8.250  -0.5757   0.09654   0.09436   0.0162   1.0000   0.0154
  -8.000  -0.5725   0.09236   0.09021   0.0133   1.0000   0.0161
  -7.750  -0.5681   0.08794   0.08582   0.0091   1.0000   0.0168
  -7.500  -0.5579   0.08348   0.08137   0.0037   1.0000   0.0172
  -6.000  -0.4718   0.02594   0.02221  -0.0389   1.0000   0.0122
  -5.750  -0.4489   0.02142   0.01709  -0.0394   1.0000   0.0122
  -5.500  -0.4239   0.02102   0.01669  -0.0395   1.0000   0.0136
  -5.250  -0.3984   0.01947   0.01491  -0.0394   1.0000   0.0147
  -5.000  -0.3727   0.01625   0.01115  -0.0390   1.0000   0.0148
  -4.750  -0.3464   0.01444   0.00901  -0.0386   1.0000   0.0157
  -4.500  -0.3198   0.01332   0.00768  -0.0381   1.0000   0.0169
  -4.250  -0.2937   0.01184   0.00597  -0.0376   1.0000   0.0183
  -4.000  -0.2678   0.01103   0.00513  -0.0372   1.0000   0.0214
  -3.750  -0.2414   0.01042   0.00444  -0.0367   1.0000   0.0241
  -3.500  -0.2153   0.00974   0.00367  -0.0362   1.0000   0.0271
  -3.250  -0.1892   0.00922   0.00313  -0.0358   1.0000   0.0339
  -3.000  -0.1631   0.00875   0.00265  -0.0353   1.0000   0.0424
  -2.750  -0.1371   0.00841   0.00233  -0.0349   1.0000   0.0549
  -2.500  -0.1111   0.00814   0.00211  -0.0345   1.0000   0.0723
  -2.250  -0.0851   0.00786   0.00190  -0.0341   1.0000   0.0947
  -2.000  -0.0591   0.00761   0.00177  -0.0338   1.0000   0.1267
  -1.750  -0.0330   0.00734   0.00167  -0.0335   1.0000   0.1727
  -1.500  -0.0042   0.00703   0.00161  -0.0339   0.9990   0.2397
  -1.250   0.0356   0.00661   0.00152  -0.0367   0.9936   0.3424
  -1.000   0.0738   0.00612   0.00146  -0.0392   0.9861   0.4710
  -0.750   0.1059   0.00508   0.00143  -0.0402   0.9760   0.7578
  -0.500   0.1442   0.00458   0.00137  -0.0416   0.9633   1.0000
  -0.250   0.1778   0.00456   0.00128  -0.0425   0.9419   1.0000
   0.000   0.2055   0.00458   0.00120  -0.0421   0.9129   1.0000
   0.250   0.2306   0.00465   0.00114  -0.0410   0.8772   1.0000
   0.500   0.2555   0.00477   0.00109  -0.0399   0.8384   1.0000
   0.750   0.2808   0.00493   0.00106  -0.0390   0.7967   1.0000
   1.000   0.3065   0.00512   0.00106  -0.0382   0.7532   1.0000
   1.250   0.3325   0.00534   0.00107  -0.0376   0.7077   1.0000
   1.500   0.3588   0.00557   0.00110  -0.0371   0.6608   1.0000
   1.750   0.3853   0.00582   0.00115  -0.0367   0.6136   1.0000
   2.000   0.4118   0.00609   0.00122  -0.0364   0.5659   1.0000
   2.250   0.4386   0.00636   0.00130  -0.0361   0.5198   1.0000
   2.500   0.4654   0.00663   0.00140  -0.0359   0.4764   1.0000
   2.750   0.4923   0.00691   0.00153  -0.0357   0.4359   1.0000
   3.000   0.5192   0.00719   0.00166  -0.0355   0.3982   1.0000
   3.250   0.5463   0.00745   0.00181  -0.0353   0.3640   1.0000
   3.500   0.5733   0.00773   0.00197  -0.0352   0.3312   1.0000
   4.000   0.6271   0.00831   0.00235  -0.0349   0.2711   1.0000
   4.250   0.6540   0.00861   0.00256  -0.0347   0.2417   1.0000
   4.500   0.6808   0.00893   0.00279  -0.0346   0.2131   1.0000
   4.750   0.7075   0.00928   0.00307  -0.0344   0.1840   1.0000
   5.000   0.7339   0.00967   0.00335  -0.0342   0.1541   1.0000
   5.250   0.7602   0.01011   0.00367  -0.0341   0.1240   1.0000
   5.500   0.7862   0.01059   0.00403  -0.0339   0.0955   1.0000
   5.750   0.8121   0.01112   0.00445  -0.0336   0.0692   1.0000
   6.000   0.8376   0.01176   0.00500  -0.0333   0.0457   1.0000
   6.250   0.8629   0.01247   0.00565  -0.0330   0.0288   1.0000
   6.500   0.8881   0.01325   0.00648  -0.0325   0.0209   1.0000
   6.750   0.9113   0.01457   0.00792  -0.0317   0.0165   1.0000
   7.000   0.9363   0.01531   0.00878  -0.0312   0.0150   1.0000
   7.250   0.9605   0.01619   0.00976  -0.0306   0.0135   1.0000
   7.500   0.9843   0.01707   0.01073  -0.0301   0.0120   1.0000
   7.750   1.0015   0.01983   0.01375  -0.0287   0.0108   1.0000
   8.000   1.0246   0.02098   0.01505  -0.0280   0.0105   1.0000
   8.250   1.0467   0.02240   0.01668  -0.0271   0.0100   1.0000
   8.500   1.0672   0.02422   0.01873  -0.0262   0.0096   1.0000
   8.750   1.0858   0.02650   0.02130  -0.0250   0.0092   1.0000
   9.000   1.1014   0.02943   0.02460  -0.0237   0.0091   1.0000
   9.250   1.1119   0.03331   0.02893  -0.0222   0.0091   1.0000
   9.500   1.1148   0.03826   0.03439  -0.0204   0.0094   1.0000
   9.750   1.1096   0.04382   0.04046  -0.0189   0.0097   1.0000
  10.000   1.0965   0.04948   0.04647  -0.0178   0.0100   1.0000
  10.250   1.0765   0.05448   0.05173  -0.0169   0.0102   1.0000
  10.500   1.0553   0.05973   0.05718  -0.0189   0.0103   1.0000
  10.750   1.0373   0.06684   0.06446  -0.0254   0.0104   1.0000
  11.000   1.0217   0.07624   0.07402  -0.0348   0.0104   1.0000
  11.250   1.0038   0.08743   0.08533  -0.0441   0.0103   1.0000
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