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AG09 (ag09-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: AG09 (ag09-il)
Reynolds number: 50,000
Max Cl/Cd: 32.35 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag09-il-50000.txt
Download as CSV file: xf-ag09-il-50000.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 =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5995   0.11592   0.10931   0.0211   1.0000   0.1507
  -8.000  -0.5788   0.10908   0.10244   0.0242   1.0000   0.1571
  -7.750  -0.5805   0.10680   0.10024   0.0211   1.0000   0.1634
  -7.500  -0.5757   0.10273   0.09624   0.0196   1.0000   0.1675
  -7.250  -0.5680   0.09923   0.09276   0.0182   1.0000   0.1768
  -7.000  -0.5595   0.09503   0.08857   0.0173   1.0000   0.1843
  -6.750  -0.5557   0.09220   0.08580   0.0101   1.0000   0.1941
  -6.500  -0.5464   0.08901   0.08261   0.0059   1.0000   0.2074
  -6.000  -0.5236   0.08049   0.07415   0.0061   1.0000   0.2365
  -5.750  -0.5112   0.07636   0.07006   0.0069   1.0000   0.2521
  -5.500  -0.4984   0.07258   0.06631   0.0074   1.0000   0.2708
  -5.250  -0.4863   0.06894   0.06268   0.0062   1.0000   0.2952
  -4.750  -0.4637   0.06220   0.05602   0.0083   1.0000   0.3638
  -4.500  -0.4560   0.05906   0.05296   0.0147   1.0000   0.4121
  -4.000  -0.4488   0.05356   0.04769   0.0340   1.0000   0.5547
  -3.750  -0.2897   0.03954   0.03127  -0.0310   1.0000   0.1765
  -3.500  -0.2514   0.03467   0.02560  -0.0337   1.0000   0.1479
  -3.250  -0.2198   0.03152   0.02184  -0.0346   1.0000   0.1496
  -3.000  -0.1881   0.02864   0.01829  -0.0349   1.0000   0.1510
  -2.750  -0.1604   0.02604   0.01556  -0.0348   1.0000   0.1586
  -2.500  -0.1316   0.02409   0.01322  -0.0345   1.0000   0.1761
  -2.250  -0.1023   0.02221   0.01102  -0.0341   1.0000   0.1937
  -2.000  -0.0755   0.02061   0.00937  -0.0334   1.0000   0.2225
  -1.750  -0.0473   0.01916   0.00792  -0.0327   1.0000   0.2637
  -1.500  -0.0199   0.01757   0.00661  -0.0320   1.0000   0.3230
  -1.250   0.0074   0.01374   0.00510  -0.0299   1.0000   1.0000
  -1.000   0.0349   0.01375   0.00440  -0.0293   1.0000   1.0000
  -0.750   0.0606   0.01376   0.00406  -0.0288   1.0000   1.0000
  -0.500   0.0859   0.01379   0.00384  -0.0283   1.0000   1.0000
  -0.250   0.1109   0.01384   0.00370  -0.0278   1.0000   1.0000
   0.000   0.1356   0.01390   0.00361  -0.0272   1.0000   1.0000
   0.250   0.1600   0.01398   0.00359  -0.0267   1.0000   1.0000
   0.500   0.1843   0.01409   0.00364  -0.0262   1.0000   1.0000
   0.750   0.2083   0.01422   0.00375  -0.0258   1.0000   1.0000
   1.000   0.2320   0.01439   0.00392  -0.0254   1.0000   1.0000
   1.250   0.2553   0.01460   0.00416  -0.0251   1.0000   1.0000
   1.500   0.2782   0.01487   0.00448  -0.0248   1.0000   1.0000
   1.750   0.3005   0.01522   0.00492  -0.0247   1.0000   1.0000
   2.000   0.3223   0.01569   0.00549  -0.0249   1.0000   1.0000
   2.250   0.3427   0.01636   0.00630  -0.0256   1.0000   1.0000
   2.500   0.3818   0.01730   0.00745  -0.0306   0.9866   1.0000
   2.750   0.4965   0.01737   0.00811  -0.0463   0.9076   1.0000
   3.000   0.5450   0.01727   0.00816  -0.0469   0.8295   1.0000
   3.250   0.5668   0.01757   0.00833  -0.0426   0.7574   1.0000
   3.500   0.5856   0.01810   0.00865  -0.0385   0.6894   1.0000
   3.750   0.6059   0.01879   0.00918  -0.0353   0.6253   1.0000
   4.000   0.6275   0.01959   0.00978  -0.0327   0.5672   1.0000
   4.250   0.6502   0.02050   0.01051  -0.0307   0.5143   1.0000
   4.500   0.6737   0.02151   0.01137  -0.0290   0.4656   1.0000
   4.750   0.6973   0.02260   0.01230  -0.0274   0.4206   1.0000
   5.000   0.7210   0.02379   0.01352  -0.0261   0.3756   1.0000
   5.250   0.7444   0.02506   0.01475  -0.0248   0.3315   1.0000
   5.500   0.7674   0.02643   0.01607  -0.0235   0.2868   1.0000
   5.750   0.7899   0.02791   0.01743  -0.0221   0.2416   1.0000
   6.000   0.8121   0.02963   0.01900  -0.0208   0.1984   1.0000
   6.250   0.8347   0.03193   0.02155  -0.0197   0.1627   1.0000
   6.500   0.8577   0.03482   0.02460  -0.0187   0.1400   1.0000
   6.750   0.8795   0.03790   0.02810  -0.0178   0.1226   1.0000
   7.000   0.9004   0.04183   0.03237  -0.0171   0.1140   1.0000
   7.250   0.9200   0.04505   0.03583  -0.0164   0.1039   1.0000
   7.500   0.9340   0.05046   0.04192  -0.0162   0.1017   1.0000
   7.750   0.9444   0.05640   0.04846  -0.0164   0.1018   1.0000
   8.000   0.9534   0.06243   0.05486  -0.0168   0.1027   1.0000
   8.250   0.9335   0.07294   0.06628  -0.0225   0.1113   1.0000
   8.500   0.9348   0.07937   0.07281  -0.0242   0.1147   1.0000
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