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AG08 (ag08-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: AG08 (ag08-il)
Reynolds number: 200,000
Max Cl/Cd: 56.62 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag08-il-200000-n5.txt
Download as CSV file: xf-ag08-il-200000-n5.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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.5921   0.10889   0.10534   0.0193   1.0000   0.0114
  -9.000  -0.5894   0.10486   0.10134   0.0178   1.0000   0.0113
  -8.750  -0.5868   0.10074   0.09724   0.0160   1.0000   0.0112
  -8.500  -0.5843   0.09621   0.09274   0.0134   1.0000   0.0114
  -8.000  -0.5761   0.09004   0.08664   0.0103   1.0000   0.0130
  -7.750  -0.5717   0.08649   0.08312   0.0079   1.0000   0.0136
  -7.500  -0.5659   0.08201   0.07868   0.0036   1.0000   0.0141
  -7.250  -0.5571   0.07663   0.07331  -0.0022   1.0000   0.0143
  -7.000  -0.5456   0.07037   0.06703  -0.0090   1.0000   0.0140
  -6.750  -0.5310   0.06278   0.05938  -0.0171   1.0000   0.0134
  -6.500  -0.5128   0.05295   0.04937  -0.0262   1.0000   0.0129
  -6.250  -0.4944   0.03536   0.03099  -0.0361   1.0000   0.0120
  -6.000  -0.4734   0.02887   0.02378  -0.0383   1.0000   0.0122
  -5.750  -0.4499   0.02522   0.01961  -0.0389   1.0000   0.0127
  -5.500  -0.4252   0.02254   0.01646  -0.0390   1.0000   0.0134
  -5.250  -0.3997   0.02031   0.01380  -0.0389   1.0000   0.0144
  -5.000  -0.3735   0.01898   0.01213  -0.0387   1.0000   0.0164
  -4.750  -0.3474   0.01736   0.01014  -0.0383   1.0000   0.0178
  -4.500  -0.3217   0.01576   0.00836  -0.0380   1.0000   0.0194
  -4.250  -0.2955   0.01483   0.00730  -0.0376   1.0000   0.0215
  -4.000  -0.2692   0.01415   0.00650  -0.0372   1.0000   0.0249
  -3.750  -0.2433   0.01323   0.00551  -0.0368   1.0000   0.0288
  -3.500  -0.2171   0.01263   0.00484  -0.0364   1.0000   0.0336
  -3.250  -0.1910   0.01209   0.00427  -0.0360   1.0000   0.0414
  -3.000  -0.1649   0.01164   0.00376  -0.0356   1.0000   0.0503
  -2.750  -0.1390   0.01124   0.00341  -0.0352   1.0000   0.0648
  -2.500  -0.1131   0.01089   0.00309  -0.0348   1.0000   0.0816
  -2.250  -0.0873   0.01057   0.00283  -0.0344   1.0000   0.1051
  -2.000  -0.0615   0.01026   0.00263  -0.0341   1.0000   0.1350
  -1.750  -0.0357   0.00995   0.00248  -0.0338   1.0000   0.1775
  -1.500  -0.0099   0.00964   0.00238  -0.0336   1.0000   0.2327
  -1.250   0.0260   0.00926   0.00227  -0.0356   0.9918   0.3161
  -1.000   0.0625   0.00881   0.00219  -0.0376   0.9795   0.4241
  -0.750   0.0958   0.00812   0.00212  -0.0388   0.9646   0.5960
  -0.500   0.1333   0.00717   0.00206  -0.0397   0.9484   1.0000
  -0.250   0.1690   0.00717   0.00194  -0.0411   0.9237   1.0000
   0.000   0.2015   0.00719   0.00185  -0.0417   0.8927   1.0000
   0.250   0.2307   0.00724   0.00176  -0.0414   0.8568   1.0000
   0.500   0.2570   0.00734   0.00170  -0.0406   0.8164   1.0000
   0.750   0.2823   0.00749   0.00165  -0.0395   0.7742   1.0000
   1.000   0.3075   0.00768   0.00164  -0.0386   0.7307   1.0000
   1.250   0.3329   0.00790   0.00165  -0.0377   0.6859   1.0000
   1.500   0.3585   0.00815   0.00168  -0.0370   0.6406   1.0000
   1.750   0.3843   0.00842   0.00175  -0.0364   0.5961   1.0000
   2.000   0.4102   0.00870   0.00183  -0.0358   0.5523   1.0000
   2.250   0.4363   0.00900   0.00194  -0.0354   0.5096   1.0000
   2.500   0.4626   0.00930   0.00207  -0.0350   0.4695   1.0000
   2.750   0.4889   0.00960   0.00225  -0.0347   0.4323   1.0000
   3.000   0.5153   0.00992   0.00242  -0.0344   0.3972   1.0000
   3.250   0.5418   0.01023   0.00262  -0.0341   0.3643   1.0000
   3.500   0.5682   0.01054   0.00284  -0.0338   0.3333   1.0000
   3.750   0.5947   0.01088   0.00311  -0.0336   0.3034   1.0000
   4.000   0.6211   0.01123   0.00338  -0.0334   0.2744   1.0000
   4.250   0.6475   0.01159   0.00367  -0.0331   0.2465   1.0000
   4.500   0.6737   0.01197   0.00399  -0.0329   0.2189   1.0000
   4.750   0.6999   0.01238   0.00437  -0.0327   0.1915   1.0000
   5.000   0.7259   0.01282   0.00475  -0.0324   0.1642   1.0000
   5.250   0.7517   0.01332   0.00518  -0.0322   0.1365   1.0000
   5.500   0.7772   0.01388   0.00566  -0.0319   0.1089   1.0000
   5.750   0.8025   0.01450   0.00620  -0.0317   0.0831   1.0000
   6.000   0.8276   0.01518   0.00685  -0.0313   0.0603   1.0000
   6.250   0.8524   0.01593   0.00759  -0.0310   0.0432   1.0000
   6.500   0.8769   0.01677   0.00844  -0.0305   0.0312   1.0000
   6.750   0.9007   0.01778   0.00950  -0.0300   0.0240   1.0000
   7.000   0.9247   0.01871   0.01060  -0.0294   0.0203   1.0000
   7.250   0.9466   0.02006   0.01202  -0.0288   0.0166   1.0000
   7.500   0.9701   0.02102   0.01320  -0.0282   0.0149   1.0000
   7.750   0.9922   0.02231   0.01468  -0.0274   0.0136   1.0000
   8.000   1.0135   0.02370   0.01624  -0.0266   0.0125   1.0000
   8.250   1.0339   0.02528   0.01798  -0.0258   0.0117   1.0000
   8.500   1.0517   0.02749   0.02040  -0.0248   0.0111   1.0000
   8.750   1.0681   0.03013   0.02334  -0.0236   0.0106   1.0000
   9.000   1.0867   0.03191   0.02546  -0.0227   0.0099   1.0000
   9.250   1.1032   0.03395   0.02784  -0.0218   0.0090   1.0000
   9.500   1.1159   0.03662   0.03092  -0.0207   0.0086   1.0000
   9.750   1.1246   0.03975   0.03444  -0.0196   0.0083   1.0000
  10.000   1.1289   0.04319   0.03828  -0.0184   0.0081   1.0000
  10.250   1.1278   0.04694   0.04239  -0.0174   0.0079   1.0000
  10.500   1.1196   0.05093   0.04671  -0.0165   0.0078   1.0000
  10.750   1.1035   0.05530   0.05135  -0.0164   0.0078   1.0000
  11.000   1.0848   0.06129   0.05760  -0.0200   0.0079   1.0000
  11.250   1.0636   0.07023   0.06678  -0.0281   0.0080   1.0000
  11.500   1.0357   0.08316   0.07993  -0.0391   0.0083   1.0000
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