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AG09 (ag09-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: AG09 (ag09-il)
Reynolds number: 100,000
Max Cl/Cd: 44.27 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag09-il-100000.txt
Download as CSV file: xf-ag09-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG09                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5869   0.11184   0.10699   0.0239   1.0000   0.0656
  -8.250  -0.5888   0.10973   0.10496   0.0184   1.0000   0.0669
  -8.000  -0.5880   0.10723   0.10252   0.0102   1.0000   0.0674
  -7.750  -0.5808   0.10220   0.09754   0.0054   1.0000   0.0680
  -7.500  -0.5683   0.09677   0.09209   0.0171   1.0000   0.0722
  -7.250  -0.5610   0.09324   0.08858   0.0147   1.0000   0.0757
  -7.000  -0.5502   0.08951   0.08487   0.0058   1.0000   0.0797
  -6.750  -0.5320   0.08488   0.08013  -0.0089   1.0000   0.0814
  -6.500  -0.5270   0.08009   0.07543  -0.0013   1.0000   0.0836
  -6.250  -0.5141   0.07637   0.07172  -0.0016   1.0000   0.0871
  -6.000  -0.4806   0.07228   0.06722  -0.0191   1.0000   0.0950
  -5.750  -0.4738   0.06686   0.06202  -0.0146   1.0000   0.0968
  -5.500  -0.4590   0.06332   0.05850  -0.0140   1.0000   0.1008
  -5.250  -0.4297   0.05851   0.05340  -0.0221   1.0000   0.1101
  -5.000  -0.4145   0.05498   0.04994  -0.0208   1.0000   0.1143
  -4.750  -0.3872   0.05077   0.04549  -0.0253   1.0000   0.1251
  -4.500  -0.3630   0.04735   0.04190  -0.0273   1.0000   0.1388
  -4.250  -0.3416   0.04416   0.03864  -0.0278   1.0000   0.1541
  -4.000  -0.3178   0.04098   0.03537  -0.0287   1.0000   0.1695
  -3.750  -0.2599   0.03083   0.02346  -0.0344   1.0000   0.0804
  -3.500  -0.2328   0.02682   0.01920  -0.0347   1.0000   0.0758
  -3.250  -0.2031   0.02391   0.01572  -0.0347   1.0000   0.0777
  -3.000  -0.1730   0.02170   0.01283  -0.0343   1.0000   0.0808
  -2.750  -0.1458   0.01936   0.01035  -0.0341   1.0000   0.0862
  -2.500  -0.1183   0.01800   0.00880  -0.0337   1.0000   0.0992
  -2.250  -0.0909   0.01662   0.00727  -0.0332   1.0000   0.1126
  -2.000  -0.0642   0.01561   0.00625  -0.0327   1.0000   0.1335
  -1.750  -0.0380   0.01463   0.00538  -0.0322   1.0000   0.1601
  -1.500  -0.0120   0.01370   0.00458  -0.0316   1.0000   0.1929
  -1.250   0.0133   0.01276   0.00396  -0.0311   1.0000   0.2468
  -1.000   0.0391   0.00974   0.00333  -0.0296   1.0000   1.0000
  -0.750   0.0652   0.00976   0.00301  -0.0291   1.0000   1.0000
  -0.500   0.0908   0.00979   0.00284  -0.0285   1.0000   1.0000
  -0.250   0.1161   0.00984   0.00275  -0.0280   1.0000   1.0000
   0.000   0.1412   0.00990   0.00270  -0.0276   1.0000   1.0000
   0.250   0.1662   0.00998   0.00271  -0.0271   1.0000   1.0000
   0.500   0.1910   0.01008   0.00277  -0.0268   1.0000   1.0000
   0.750   0.2156   0.01021   0.00289  -0.0264   1.0000   1.0000
   1.000   0.2399   0.01039   0.00307  -0.0261   1.0000   1.0000
   1.250   0.2640   0.01062   0.00333  -0.0260   1.0000   1.0000
   1.500   0.2878   0.01093   0.00368  -0.0261   1.0000   1.0000
   1.750   0.3562   0.01100   0.00390  -0.0345   0.9701   1.0000
   2.000   0.4111   0.01094   0.00395  -0.0392   0.9280   1.0000
   2.250   0.4486   0.01093   0.00396  -0.0398   0.8730   1.0000
   2.500   0.4736   0.01105   0.00399  -0.0376   0.8106   1.0000
   2.750   0.4953   0.01131   0.00409  -0.0349   0.7443   1.0000
   3.000   0.5173   0.01170   0.00422  -0.0326   0.6746   1.0000
   3.250   0.5401   0.01220   0.00442  -0.0308   0.6058   1.0000
   3.500   0.5637   0.01277   0.00470  -0.0294   0.5421   1.0000
   3.750   0.5878   0.01339   0.00510  -0.0283   0.4854   1.0000
   4.000   0.6124   0.01405   0.00552  -0.0275   0.4360   1.0000
   4.250   0.6373   0.01474   0.00601  -0.0268   0.3917   1.0000
   4.500   0.6623   0.01543   0.00656  -0.0262   0.3512   1.0000
   4.750   0.6875   0.01610   0.00715  -0.0256   0.3122   1.0000
   5.000   0.7126   0.01682   0.00784  -0.0251   0.2751   1.0000
   5.250   0.7372   0.01759   0.00847  -0.0245   0.2386   1.0000
   5.500   0.7616   0.01837   0.00920  -0.0240   0.1984   1.0000
   5.750   0.7856   0.01936   0.01011  -0.0235   0.1538   1.0000
   6.000   0.8091   0.02074   0.01139  -0.0228   0.1108   1.0000
   6.250   0.8325   0.02256   0.01319  -0.0220   0.0836   1.0000
   6.500   0.8562   0.02447   0.01511  -0.0212   0.0681   1.0000
   6.750   0.8802   0.02693   0.01776  -0.0203   0.0595   1.0000
   7.000   0.9023   0.02991   0.02072  -0.0197   0.0526   1.0000
   7.250   0.9257   0.03282   0.02428  -0.0186   0.0506   1.0000
   7.500   0.9458   0.03669   0.02881  -0.0175   0.0496   1.0000
   7.750   0.9623   0.04104   0.03387  -0.0166   0.0488   1.0000
   8.000   0.9748   0.04570   0.03918  -0.0160   0.0477   1.0000
   8.250   0.9820   0.05114   0.04520  -0.0157   0.0475   1.0000
   8.500   0.9830   0.05744   0.05200  -0.0161   0.0486   1.0000
   8.750   0.9808   0.06367   0.05855  -0.0169   0.0502   1.0000
   9.000   0.9793   0.06956   0.06457  -0.0176   0.0517   1.0000
   9.250   0.9283   0.08652   0.08205  -0.0307   0.0675   1.0000
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