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AG09 (ag09-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: AG09 (ag09-il)
Reynolds number: 100,000
Max Cl/Cd: 44.1 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag09-il-100000-n5.txt
Download as CSV file: xf-ag09-il-100000-n5.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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4663   0.09016   0.08559   0.0094   1.0000   0.0487
  -7.750  -0.4657   0.08627   0.08174   0.0080   1.0000   0.0496
  -7.250  -0.5606   0.09033   0.08563   0.0115   1.0000   0.0456
  -7.000  -0.5506   0.08644   0.08175   0.0084   1.0000   0.0472
  -6.750  -0.5388   0.08207   0.07739   0.0041   1.0000   0.0479
  -6.250  -0.5015   0.06767   0.06277  -0.0118   1.0000   0.0297
  -6.000  -0.4849   0.06297   0.05801  -0.0149   1.0000   0.0287
  -5.750  -0.4645   0.05779   0.05271  -0.0189   1.0000   0.0278
  -5.500  -0.4414   0.05230   0.04701  -0.0229   1.0000   0.0271
  -5.250  -0.4162   0.04664   0.04108  -0.0266   1.0000   0.0265
  -5.000  -0.3895   0.04088   0.03496  -0.0297   1.0000   0.0262
  -4.750  -0.3620   0.03548   0.02910  -0.0319   1.0000   0.0265
  -4.500  -0.3337   0.03071   0.02371  -0.0333   1.0000   0.0287
  -4.250  -0.3049   0.02662   0.01886  -0.0340   1.0000   0.0302
  -4.000  -0.2770   0.02342   0.01502  -0.0341   1.0000   0.0311
  -3.750  -0.2503   0.02117   0.01248  -0.0342   1.0000   0.0329
  -3.500  -0.2233   0.01991   0.01100  -0.0340   1.0000   0.0370
  -3.250  -0.1956   0.01850   0.00916  -0.0335   1.0000   0.0417
  -3.000  -0.1691   0.01708   0.00770  -0.0332   1.0000   0.0460
  -2.750  -0.1422   0.01627   0.00673  -0.0328   1.0000   0.0550
  -2.500  -0.1159   0.01537   0.00585  -0.0325   1.0000   0.0639
  -2.250  -0.0895   0.01474   0.00518  -0.0321   1.0000   0.0784
  -2.000  -0.0633   0.01413   0.00456  -0.0317   1.0000   0.0943
  -1.750  -0.0373   0.01359   0.00413  -0.0313   1.0000   0.1144
  -1.500  -0.0113   0.01313   0.00374  -0.0309   1.0000   0.1403
  -1.250   0.0147   0.01268   0.00340  -0.0306   1.0000   0.1747
  -1.000   0.0404   0.01217   0.00317  -0.0303   1.0000   0.2393
  -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   0.2100   0.01005   0.00271  -0.0305   0.9754   1.0000
   0.750   0.2519   0.01012   0.00272  -0.0333   0.9446   1.0000
   1.000   0.2906   0.01018   0.00272  -0.0351   0.9060   1.0000
   1.250   0.3246   0.01027   0.00272  -0.0358   0.8601   1.0000
   1.500   0.3530   0.01041   0.00273  -0.0351   0.8092   1.0000
   1.750   0.3784   0.01062   0.00278  -0.0338   0.7563   1.0000
   2.000   0.4030   0.01088   0.00284  -0.0324   0.7019   1.0000
   2.250   0.4278   0.01118   0.00293  -0.0313   0.6464   1.0000
   2.500   0.4528   0.01152   0.00306  -0.0303   0.5911   1.0000
   2.750   0.4781   0.01190   0.00326  -0.0295   0.5377   1.0000
   3.000   0.5036   0.01231   0.00347  -0.0288   0.4880   1.0000
   3.250   0.5292   0.01273   0.00372  -0.0282   0.4426   1.0000
   3.500   0.5549   0.01316   0.00401  -0.0277   0.4015   1.0000
   3.750   0.5807   0.01362   0.00439  -0.0273   0.3640   1.0000
   4.000   0.6067   0.01407   0.00476  -0.0269   0.3293   1.0000
   4.250   0.6326   0.01455   0.00517  -0.0266   0.2972   1.0000
   4.500   0.6584   0.01505   0.00561  -0.0262   0.2668   1.0000
   4.750   0.6842   0.01556   0.00610  -0.0259   0.2369   1.0000
   5.000   0.7100   0.01610   0.00668  -0.0256   0.2075   1.0000
   5.250   0.7356   0.01669   0.00725  -0.0253   0.1773   1.0000
   5.500   0.7609   0.01735   0.00788  -0.0250   0.1456   1.0000
   5.750   0.7859   0.01813   0.00860  -0.0247   0.1123   1.0000
   6.000   0.8103   0.01912   0.00955  -0.0244   0.0803   1.0000
   6.250   0.8341   0.02036   0.01075  -0.0240   0.0569   1.0000
   6.500   0.8573   0.02173   0.01214  -0.0235   0.0428   1.0000
   6.750   0.8801   0.02327   0.01385  -0.0227   0.0355   1.0000
   7.000   0.9015   0.02501   0.01565  -0.0221   0.0301   1.0000
   7.250   0.9243   0.02655   0.01749  -0.0213   0.0263   1.0000
   7.500   0.9457   0.02848   0.01968  -0.0205   0.0241   1.0000
   7.750   0.9659   0.03068   0.02208  -0.0198   0.0226   1.0000
   8.000   0.9837   0.03379   0.02545  -0.0190   0.0212   1.0000
   8.250   1.0028   0.03628   0.02847  -0.0182   0.0197   1.0000
   8.500   1.0189   0.03928   0.03196  -0.0174   0.0183   1.0000
   8.750   1.0308   0.04303   0.03623  -0.0167   0.0176   1.0000
   9.000   1.0378   0.04739   0.04112  -0.0162   0.0173   1.0000
   9.250   1.0385   0.05228   0.04652  -0.0160   0.0172   1.0000
   9.500   1.0321   0.05769   0.05237  -0.0165   0.0171   1.0000
   9.750   1.0176   0.06349   0.05852  -0.0181   0.0172   1.0000
  10.000   0.9962   0.07010   0.06535  -0.0224   0.0174   1.0000
  10.250   0.9724   0.08104   0.07645  -0.0333   0.0179   1.0000
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