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AG10 (ag10-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: AG10 (ag10-il)
Reynolds number: 500,000
Max Cl/Cd: 69.79 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag10-il-500000.txt
Download as CSV file: xf-ag10-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG10                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.6393   0.12046   0.11816   0.0356   1.0000   0.0148
  -9.250  -0.6366   0.11627   0.11399   0.0327   1.0000   0.0149
  -9.000  -0.6329   0.11213   0.10986   0.0304   1.0000   0.0150
  -8.750  -0.6314   0.10760   0.10535   0.0298   1.0000   0.0152
  -8.500  -0.6241   0.10455   0.10230   0.0301   1.0000   0.0154
  -8.250  -0.6184   0.10135   0.09911   0.0291   1.0000   0.0157
  -8.000  -0.6135   0.09798   0.09576   0.0276   1.0000   0.0160
  -7.750  -0.6088   0.09454   0.09234   0.0257   1.0000   0.0164
  -7.500  -0.6040   0.09100   0.08881   0.0232   1.0000   0.0169
  -7.250  -0.5949   0.08674   0.08456   0.0188   1.0000   0.0177
  -6.250  -0.5250   0.06223   0.05988  -0.0102   1.0000   0.0197
  -6.000  -0.5076   0.05892   0.05653  -0.0122   1.0000   0.0200
  -5.750  -0.4869   0.05533   0.05288  -0.0148   1.0000   0.0206
  -5.500  -0.4632   0.05118   0.04864  -0.0181   1.0000   0.0214
  -5.250  -0.4355   0.04596   0.04325  -0.0221   1.0000   0.0229
  -5.000  -0.3997   0.03479   0.03146  -0.0281   1.0000   0.0251
  -4.750  -0.3729   0.02396   0.01981  -0.0305   1.0000   0.0199
  -4.500  -0.3452   0.01964   0.01501  -0.0310   1.0000   0.0179
  -4.250  -0.3166   0.01673   0.01155  -0.0311   1.0000   0.0185
  -4.000  -0.2882   0.01466   0.00906  -0.0310   1.0000   0.0189
  -3.750  -0.2599   0.01286   0.00693  -0.0309   1.0000   0.0195
  -3.500  -0.2323   0.01157   0.00556  -0.0308   1.0000   0.0208
  -3.250  -0.2044   0.01095   0.00488  -0.0306   1.0000   0.0225
  -3.000  -0.1764   0.01059   0.00444  -0.0304   1.0000   0.0245
  -2.750  -0.1487   0.00975   0.00361  -0.0304   1.0000   0.0281
  -2.500  -0.1208   0.00938   0.00320  -0.0302   1.0000   0.0311
  -2.250  -0.0931   0.00882   0.00266  -0.0300   1.0000   0.0367
  -2.000  -0.0652   0.00839   0.00223  -0.0299   1.0000   0.0431
  -1.750  -0.0374   0.00811   0.00198  -0.0297   1.0000   0.0530
  -1.500  -0.0098   0.00780   0.00176  -0.0296   1.0000   0.0689
  -1.250   0.0177   0.00757   0.00160  -0.0294   1.0000   0.0908
  -1.000   0.0449   0.00731   0.00149  -0.0292   1.0000   0.1228
  -0.750   0.0720   0.00703   0.00142  -0.0291   1.0000   0.1758
  -0.500   0.0988   0.00662   0.00140  -0.0290   1.0000   0.2851
  -0.250   0.1163   0.00469   0.00145  -0.0270   1.0000   0.8812
   0.000   0.1554   0.00453   0.00135  -0.0291   0.9866   1.0000
   0.250   0.1959   0.00454   0.00129  -0.0316   0.9537   1.0000
   0.500   0.2244   0.00463   0.00124  -0.0311   0.8976   1.0000
   0.750   0.2469   0.00485   0.00120  -0.0293   0.8279   1.0000
   1.000   0.2711   0.00518   0.00118  -0.0280   0.7492   1.0000
   1.250   0.2966   0.00556   0.00119  -0.0273   0.6629   1.0000
   1.500   0.3231   0.00595   0.00123  -0.0269   0.5819   1.0000
   1.750   0.3502   0.00631   0.00129  -0.0267   0.5125   1.0000
   2.000   0.3776   0.00663   0.00137  -0.0266   0.4573   1.0000
   2.250   0.4052   0.00691   0.00147  -0.0265   0.4130   1.0000
   2.500   0.4328   0.00718   0.00158  -0.0264   0.3758   1.0000
   2.750   0.4605   0.00743   0.00169  -0.0263   0.3425   1.0000
   3.000   0.4882   0.00767   0.00182  -0.0263   0.3114   1.0000
   3.250   0.5159   0.00792   0.00197  -0.0262   0.2819   1.0000
   3.500   0.5435   0.00818   0.00213  -0.0262   0.2531   1.0000
   3.750   0.5710   0.00846   0.00230  -0.0261   0.2249   1.0000
   4.000   0.5985   0.00876   0.00249  -0.0260   0.1971   1.0000
   4.250   0.6259   0.00908   0.00272  -0.0260   0.1703   1.0000
   4.500   0.6532   0.00940   0.00295  -0.0259   0.1449   1.0000
   4.750   0.6805   0.00975   0.00321  -0.0258   0.1210   1.0000
   5.000   0.7076   0.01014   0.00351  -0.0257   0.0981   1.0000
   5.250   0.7345   0.01059   0.00387  -0.0256   0.0756   1.0000
   5.500   0.7613   0.01105   0.00426  -0.0255   0.0556   1.0000
   5.750   0.7880   0.01165   0.00479  -0.0253   0.0392   1.0000
   6.000   0.8144   0.01230   0.00542  -0.0251   0.0287   1.0000
   6.250   0.8408   0.01298   0.00613  -0.0249   0.0228   1.0000
   6.500   0.8664   0.01390   0.00712  -0.0246   0.0195   1.0000
   6.750   0.8925   0.01458   0.00787  -0.0242   0.0176   1.0000
   7.000   0.9173   0.01561   0.00897  -0.0239   0.0160   1.0000
   7.250   0.9414   0.01689   0.01039  -0.0233   0.0151   1.0000
   7.500   0.9665   0.01772   0.01133  -0.0229   0.0141   1.0000
   7.750   0.9913   0.01858   0.01228  -0.0225   0.0131   1.0000
   8.000   1.0148   0.01976   0.01356  -0.0221   0.0126   1.0000
   8.250   1.0363   0.02165   0.01559  -0.0214   0.0121   1.0000
   8.500   1.0556   0.02438   0.01861  -0.0205   0.0118   1.0000
   8.750   1.0770   0.02610   0.02063  -0.0198   0.0116   1.0000
   9.000   1.0960   0.02846   0.02335  -0.0190   0.0114   1.0000
   9.250   1.1114   0.03155   0.02686  -0.0181   0.0113   1.0000
   9.500   1.1216   0.03554   0.03132  -0.0171   0.0113   1.0000
   9.750   1.1252   0.04040   0.03664  -0.0162   0.0115   1.0000
  10.000   1.1206   0.04602   0.04268  -0.0157   0.0117   1.0000
  10.250   1.1200   0.05035   0.04719  -0.0152   0.0120   1.0000
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