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AG08 (ag08-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: AG08 (ag08-il)
Reynolds number: 200,000
Max Cl/Cd: 58.34 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag08-il-200000.txt
Download as CSV file: xf-ag08-il-200000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4578   0.08433   0.08113   0.0029   1.0000   0.0432
  -7.750  -0.4582   0.08026   0.07707   0.0012   1.0000   0.0447
  -7.500  -0.5591   0.08709   0.08381   0.0042   1.0000   0.0389
  -7.250  -0.5493   0.08410   0.08083   0.0035   1.0000   0.0413
  -7.000  -0.5388   0.07982   0.07655  -0.0009   1.0000   0.0435
  -6.750  -0.5244   0.07461   0.07133  -0.0080   1.0000   0.0460
  -6.500  -0.4917   0.06674   0.06317  -0.0253   1.0000   0.0486
  -6.250  -0.4830   0.05924   0.05563  -0.0285   1.0000   0.0497
  -6.000  -0.4704   0.05622   0.05263  -0.0276   1.0000   0.0511
  -5.750  -0.4529   0.05295   0.04932  -0.0286   1.0000   0.0531
  -5.500  -0.4307   0.04855   0.04477  -0.0316   1.0000   0.0568
  -5.250  -0.4032   0.04137   0.03701  -0.0372   1.0000   0.0632
  -5.000  -0.3750   0.03001   0.02477  -0.0388   1.0000   0.0351
  -4.750  -0.3502   0.02521   0.01939  -0.0392   1.0000   0.0341
  -4.500  -0.3228   0.02314   0.01675  -0.0388   1.0000   0.0364
  -4.250  -0.2963   0.02025   0.01334  -0.0384   1.0000   0.0366
  -4.000  -0.2703   0.01680   0.00945  -0.0381   1.0000   0.0386
  -3.750  -0.2443   0.01580   0.00838  -0.0377   1.0000   0.0440
  -3.500  -0.2173   0.01496   0.00727  -0.0370   1.0000   0.0488
  -3.250  -0.1917   0.01342   0.00576  -0.0365   1.0000   0.0570
  -3.000  -0.1658   0.01257   0.00491  -0.0360   1.0000   0.0690
  -2.750  -0.1398   0.01187   0.00421  -0.0355   1.0000   0.0857
  -2.500  -0.1139   0.01117   0.00361  -0.0351   1.0000   0.1094
  -2.250  -0.0880   0.01061   0.00318  -0.0347   1.0000   0.1458
  -2.000  -0.0622   0.01000   0.00286  -0.0345   1.0000   0.2066
  -1.750  -0.0365   0.00934   0.00270  -0.0344   1.0000   0.3237
  -1.500  -0.0115   0.00855   0.00260  -0.0341   1.0000   0.4974
  -1.250   0.0148   0.00715   0.00251  -0.0326   1.0000   1.0000
  -1.000   0.0404   0.00719   0.00240  -0.0321   1.0000   1.0000
  -0.750   0.0659   0.00724   0.00234  -0.0317   1.0000   1.0000
  -0.500   0.0915   0.00731   0.00232  -0.0314   1.0000   1.0000
  -0.250   0.1170   0.00740   0.00233  -0.0311   1.0000   1.0000
   0.000   0.1423   0.00752   0.00239  -0.0308   1.0000   1.0000
   0.250   0.1785   0.00762   0.00245  -0.0328   0.9947   1.0000
   0.500   0.2285   0.00763   0.00244  -0.0375   0.9804   1.0000
   0.750   0.2738   0.00759   0.00240  -0.0409   0.9616   1.0000
   1.000   0.3146   0.00755   0.00234  -0.0432   0.9376   1.0000
   1.250   0.3484   0.00752   0.00229  -0.0438   0.9057   1.0000
   1.500   0.3762   0.00754   0.00226  -0.0430   0.8671   1.0000
   1.750   0.4008   0.00764   0.00224  -0.0415   0.8229   1.0000
   2.000   0.4246   0.00783   0.00225  -0.0399   0.7747   1.0000
   2.250   0.4486   0.00807   0.00229  -0.0385   0.7223   1.0000
   2.500   0.4730   0.00838   0.00237  -0.0373   0.6683   1.0000
   2.750   0.4978   0.00872   0.00251  -0.0364   0.6136   1.0000
   3.000   0.5228   0.00910   0.00266  -0.0356   0.5597   1.0000
   3.250   0.5480   0.00951   0.00284  -0.0349   0.5089   1.0000
   3.500   0.5736   0.00990   0.00306  -0.0344   0.4616   1.0000
   3.750   0.5993   0.01032   0.00334  -0.0339   0.4183   1.0000
   4.000   0.6250   0.01074   0.00362  -0.0335   0.3783   1.0000
   4.250   0.6508   0.01118   0.00393  -0.0332   0.3412   1.0000
   4.500   0.6767   0.01160   0.00426  -0.0328   0.3050   1.0000
   4.750   0.7025   0.01206   0.00465  -0.0325   0.2704   1.0000
   5.000   0.7283   0.01252   0.00505  -0.0322   0.2352   1.0000
   5.250   0.7538   0.01305   0.00548  -0.0319   0.1997   1.0000
   5.500   0.7790   0.01366   0.00597  -0.0316   0.1630   1.0000
   5.750   0.8040   0.01436   0.00655  -0.0313   0.1248   1.0000
   6.000   0.8283   0.01528   0.00737  -0.0308   0.0879   1.0000
   6.250   0.8514   0.01657   0.00852  -0.0302   0.0600   1.0000
   6.500   0.8741   0.01802   0.00997  -0.0294   0.0443   1.0000
   6.750   0.8969   0.01945   0.01146  -0.0285   0.0366   1.0000
   7.000   0.9192   0.02117   0.01333  -0.0275   0.0317   1.0000
   7.250   0.9424   0.02253   0.01481  -0.0267   0.0280   1.0000
   7.500   0.9634   0.02493   0.01733  -0.0257   0.0259   1.0000
   7.750   0.9815   0.02931   0.02213  -0.0245   0.0248   1.0000
   8.000   1.0012   0.03210   0.02536  -0.0233   0.0244   1.0000
   8.250   1.0181   0.03537   0.02912  -0.0221   0.0241   1.0000
   8.500   1.0340   0.03804   0.03227  -0.0209   0.0230   1.0000
   8.750   1.0449   0.04173   0.03647  -0.0197   0.0222   1.0000
   9.000   1.0476   0.04692   0.04218  -0.0185   0.0225   1.0000
   9.250   1.0439   0.05246   0.04816  -0.0176   0.0230   1.0000
   9.500   1.0342   0.05806   0.05410  -0.0173   0.0235   1.0000
   9.750   1.0187   0.06350   0.05979  -0.0175   0.0240   1.0000
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