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AG08 (ag08-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: AG08 (ag08-il)
Reynolds number: 100,000
Max Cl/Cd: 45.12 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag08-il-100000-n5.txt
Download as CSV file: xf-ag08-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5679   0.10455   0.09969   0.0147   1.0000   0.0406
  -8.250  -0.5703   0.09979   0.09499   0.0064   1.0000   0.0363
  -8.000  -0.5669   0.09296   0.08816   0.0061   1.0000   0.0253
  -7.750  -0.5630   0.08904   0.08428   0.0045   1.0000   0.0247
  -7.500  -0.5569   0.08462   0.07989   0.0012   1.0000   0.0242
  -7.250  -0.5488   0.07978   0.07506  -0.0031   1.0000   0.0238
  -7.000  -0.5385   0.07448   0.06977  -0.0082   1.0000   0.0233
  -6.750  -0.5256   0.06873   0.06398  -0.0139   1.0000   0.0229
  -6.500  -0.5100   0.06247   0.05764  -0.0199   1.0000   0.0226
  -6.250  -0.4914   0.05565   0.05066  -0.0258   1.0000   0.0228
  -6.000  -0.4705   0.04835   0.04307  -0.0311   1.0000   0.0237
  -5.750  -0.4483   0.04114   0.03536  -0.0351   1.0000   0.0246
  -5.500  -0.4252   0.03518   0.02874  -0.0373   1.0000   0.0251
  -5.250  -0.4009   0.03044   0.02329  -0.0383   1.0000   0.0255
  -5.000  -0.3759   0.02659   0.01878  -0.0387   1.0000   0.0263
  -4.750  -0.3510   0.02400   0.01581  -0.0388   1.0000   0.0280
  -4.500  -0.3253   0.02269   0.01425  -0.0386   1.0000   0.0321
  -4.250  -0.2985   0.02076   0.01187  -0.0381   1.0000   0.0354
  -4.000  -0.2724   0.01892   0.00979  -0.0377   1.0000   0.0383
  -3.750  -0.2465   0.01803   0.00881  -0.0373   1.0000   0.0453
  -3.500  -0.2206   0.01682   0.00746  -0.0368   1.0000   0.0516
  -3.250  -0.1947   0.01604   0.00658  -0.0363   1.0000   0.0622
  -3.000  -0.1689   0.01525   0.00577  -0.0358   1.0000   0.0751
  -2.750  -0.1433   0.01457   0.00517  -0.0355   1.0000   0.0946
  -2.500  -0.1176   0.01397   0.00462  -0.0351   1.0000   0.1189
  -2.250  -0.0921   0.01343   0.00418  -0.0347   1.0000   0.1564
  -2.000  -0.0666   0.01289   0.00387  -0.0343   1.0000   0.2099
  -1.750  -0.0414   0.01237   0.00364  -0.0340   1.0000   0.2884
  -1.500  -0.0162   0.01183   0.00348  -0.0336   1.0000   0.3981
  -1.250   0.0076   0.01102   0.00332  -0.0328   1.0000   0.5759
  -1.000   0.0355   0.01001   0.00318  -0.0316   1.0000   1.0000
  -0.750   0.0606   0.01006   0.00307  -0.0311   1.0000   1.0000
  -0.500   0.0854   0.01012   0.00301  -0.0307   1.0000   1.0000
  -0.250   0.1102   0.01021   0.00299  -0.0302   1.0000   1.0000
   0.000   0.1465   0.01029   0.00297  -0.0322   0.9901   1.0000
   0.250   0.1901   0.01036   0.00296  -0.0354   0.9716   1.0000
   0.500   0.2304   0.01041   0.00296  -0.0379   0.9486   1.0000
   0.750   0.2690   0.01045   0.00295  -0.0398   0.9216   1.0000
   1.000   0.3049   0.01049   0.00294  -0.0410   0.8892   1.0000
   1.250   0.3371   0.01055   0.00293  -0.0413   0.8512   1.0000
   1.500   0.3658   0.01065   0.00295  -0.0408   0.8087   1.0000
   1.750   0.3922   0.01081   0.00297  -0.0397   0.7631   1.0000
   2.000   0.4176   0.01102   0.00303  -0.0386   0.7155   1.0000
   2.250   0.4428   0.01128   0.00311  -0.0375   0.6666   1.0000
   2.500   0.4681   0.01159   0.00324  -0.0365   0.6179   1.0000
   2.750   0.4934   0.01193   0.00344  -0.0357   0.5700   1.0000
   3.000   0.5186   0.01231   0.00364  -0.0349   0.5239   1.0000
   3.250   0.5439   0.01270   0.00388  -0.0342   0.4803   1.0000
   3.500   0.5693   0.01312   0.00416  -0.0336   0.4395   1.0000
   3.750   0.5948   0.01354   0.00452  -0.0331   0.4010   1.0000
   4.000   0.6203   0.01399   0.00487  -0.0326   0.3650   1.0000
   4.250   0.6458   0.01446   0.00526  -0.0321   0.3306   1.0000
   4.500   0.6712   0.01494   0.00569  -0.0317   0.2976   1.0000
   4.750   0.6967   0.01544   0.00615  -0.0313   0.2647   1.0000
   5.000   0.7219   0.01600   0.00670  -0.0309   0.2328   1.0000
   5.250   0.7470   0.01659   0.00726  -0.0305   0.2010   1.0000
   5.500   0.7719   0.01723   0.00788  -0.0301   0.1684   1.0000
   5.750   0.7964   0.01800   0.00857  -0.0298   0.1360   1.0000
   6.000   0.8206   0.01886   0.00944  -0.0293   0.1048   1.0000
   6.250   0.8442   0.01991   0.01043  -0.0289   0.0783   1.0000
   6.500   0.8668   0.02118   0.01165  -0.0283   0.0586   1.0000
   6.750   0.8892   0.02251   0.01303  -0.0276   0.0449   1.0000
   7.000   0.9113   0.02393   0.01463  -0.0268   0.0373   1.0000
   7.250   0.9324   0.02553   0.01639  -0.0260   0.0314   1.0000
   7.500   0.9536   0.02712   0.01816  -0.0251   0.0275   1.0000
   7.750   0.9722   0.02939   0.02061  -0.0241   0.0252   1.0000
   8.000   0.9920   0.03177   0.02330  -0.0230   0.0236   1.0000
   8.250   1.0115   0.03392   0.02585  -0.0221   0.0212   1.0000
   8.500   1.0291   0.03616   0.02842  -0.0212   0.0193   1.0000
   8.750   1.0440   0.03900   0.03161  -0.0203   0.0184   1.0000
   9.000   1.0558   0.04220   0.03518  -0.0193   0.0177   1.0000
   9.250   1.0632   0.04596   0.03931  -0.0184   0.0173   1.0000
   9.500   1.0650   0.05029   0.04406  -0.0175   0.0170   1.0000
   9.750   1.0608   0.05495   0.04912  -0.0168   0.0168   1.0000
  10.000   1.0513   0.05962   0.05415  -0.0166   0.0168   1.0000
  10.250   1.0363   0.06401   0.05880  -0.0167   0.0168   1.0000
  10.500   1.0195   0.06926   0.06429  -0.0195   0.0169   1.0000
  10.750   1.0010   0.07671   0.07197  -0.0261   0.0171   1.0000
  11.000   0.9652   0.09255   0.08806  -0.0403   0.0180   1.0000
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