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AG09 (ag09-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: AG09 (ag09-il)
Reynolds number: 200,000
Max Cl/Cd: 55.62 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag09-il-200000-n5.txt
Download as CSV file: xf-ag09-il-200000-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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5980   0.11062   0.10714   0.0221   1.0000   0.0215
  -8.500  -0.5932   0.10665   0.10319   0.0201   1.0000   0.0216
  -8.250  -0.5885   0.10264   0.09920   0.0179   1.0000   0.0216
  -8.000  -0.5837   0.09859   0.09518   0.0155   1.0000   0.0216
  -7.750  -0.5775   0.09433   0.09094   0.0124   1.0000   0.0216
  -7.500  -0.5684   0.08969   0.08631   0.0084   1.0000   0.0215
  -7.250  -0.5679   0.08592   0.08256   0.0124   1.0000   0.0191
  -7.000  -0.5548   0.07678   0.07338   0.0012   1.0000   0.0146
  -6.750  -0.5425   0.07184   0.06842  -0.0031   1.0000   0.0143
  -6.500  -0.5261   0.06643   0.06296  -0.0084   1.0000   0.0141
  -6.250  -0.5068   0.06062   0.05707  -0.0138   1.0000   0.0140
  -6.000  -0.4848   0.05441   0.05072  -0.0192   1.0000   0.0139
  -5.750  -0.4581   0.04665   0.04271  -0.0250   1.0000   0.0146
  -5.500  -0.4338   0.03979   0.03554  -0.0290   1.0000   0.0154
  -5.250  -0.4074   0.03118   0.02633  -0.0322   1.0000   0.0154
  -5.000  -0.3810   0.02500   0.01938  -0.0337   1.0000   0.0154
  -4.750  -0.3545   0.02170   0.01554  -0.0341   1.0000   0.0159
  -4.500  -0.3276   0.01949   0.01288  -0.0342   1.0000   0.0166
  -4.250  -0.3005   0.01769   0.01073  -0.0341   1.0000   0.0176
  -4.000  -0.2733   0.01626   0.00900  -0.0339   1.0000   0.0191
  -3.750  -0.2460   0.01544   0.00794  -0.0336   1.0000   0.0215
  -3.500  -0.2193   0.01403   0.00639  -0.0333   1.0000   0.0237
  -3.250  -0.1923   0.01325   0.00552  -0.0330   1.0000   0.0263
  -3.000  -0.1653   0.01267   0.00483  -0.0327   1.0000   0.0304
  -2.750  -0.1386   0.01200   0.00417  -0.0325   1.0000   0.0364
  -2.500  -0.1117   0.01150   0.00361  -0.0322   1.0000   0.0434
  -2.250  -0.0849   0.01114   0.00324  -0.0319   1.0000   0.0542
  -2.000  -0.0583   0.01078   0.00291  -0.0316   1.0000   0.0672
  -1.750  -0.0318   0.01047   0.00267  -0.0313   1.0000   0.0822
  -1.500  -0.0055   0.01020   0.00248  -0.0310   1.0000   0.1009
  -1.250   0.0208   0.00996   0.00229  -0.0307   1.0000   0.1215
  -1.000   0.0468   0.00971   0.00217  -0.0304   1.0000   0.1512
  -0.750   0.0786   0.00941   0.00206  -0.0314   0.9922   0.1999
  -0.500   0.1158   0.00885   0.00197  -0.0337   0.9718   0.3321
  -0.250   0.1488   0.00700   0.00191  -0.0341   0.9475   1.0000
   0.000   0.1848   0.00703   0.00181  -0.0355   0.9119   1.0000
   0.250   0.2152   0.00709   0.00171  -0.0355   0.8684   1.0000
   0.500   0.2412   0.00721   0.00165  -0.0345   0.8213   1.0000
   0.750   0.2659   0.00739   0.00160  -0.0333   0.7735   1.0000
   1.000   0.2908   0.00760   0.00158  -0.0322   0.7246   1.0000
   1.250   0.3161   0.00785   0.00158  -0.0313   0.6752   1.0000
   1.500   0.3418   0.00812   0.00161  -0.0306   0.6247   1.0000
   1.750   0.3678   0.00841   0.00167  -0.0300   0.5745   1.0000
   2.000   0.3940   0.00871   0.00175  -0.0296   0.5259   1.0000
   2.250   0.4204   0.00902   0.00185  -0.0292   0.4799   1.0000
   2.500   0.4469   0.00933   0.00198  -0.0289   0.4381   1.0000
   2.750   0.4736   0.00964   0.00216  -0.0287   0.4003   1.0000
   3.000   0.5003   0.00996   0.00233  -0.0285   0.3652   1.0000
   3.250   0.5272   0.01027   0.00253  -0.0283   0.3334   1.0000
   3.500   0.5540   0.01059   0.00276  -0.0281   0.3036   1.0000
   3.750   0.5808   0.01092   0.00303  -0.0279   0.2752   1.0000
   4.000   0.6075   0.01126   0.00330  -0.0277   0.2485   1.0000
   4.250   0.6342   0.01162   0.00360  -0.0275   0.2226   1.0000
   4.500   0.6607   0.01200   0.00392  -0.0274   0.1971   1.0000
   4.750   0.6872   0.01240   0.00427  -0.0272   0.1716   1.0000
   5.000   0.7136   0.01283   0.00469  -0.0270   0.1457   1.0000
   5.250   0.7398   0.01333   0.00512  -0.0268   0.1186   1.0000
   5.500   0.7656   0.01392   0.00561  -0.0267   0.0901   1.0000
   5.750   0.7911   0.01459   0.00619  -0.0264   0.0632   1.0000
   6.000   0.8162   0.01543   0.00696  -0.0262   0.0413   1.0000
   6.250   0.8413   0.01630   0.00786  -0.0258   0.0293   1.0000
   6.500   0.8660   0.01725   0.00888  -0.0254   0.0221   1.0000
   6.750   0.8905   0.01828   0.01008  -0.0249   0.0190   1.0000
   7.000   0.9148   0.01924   0.01116  -0.0245   0.0163   1.0000
   7.250   0.9374   0.02059   0.01267  -0.0240   0.0140   1.0000
   7.500   0.9605   0.02185   0.01413  -0.0234   0.0129   1.0000
   7.750   0.9827   0.02333   0.01581  -0.0227   0.0120   1.0000
   8.000   1.0042   0.02495   0.01763  -0.0220   0.0113   1.0000
   8.250   1.0250   0.02672   0.01962  -0.0212   0.0108   1.0000
   8.500   1.0446   0.02870   0.02185  -0.0205   0.0104   1.0000
   8.750   1.0623   0.03107   0.02450  -0.0197   0.0101   1.0000
   9.000   1.0772   0.03401   0.02783  -0.0189   0.0098   1.0000
   9.250   1.0882   0.03762   0.03186  -0.0180   0.0097   1.0000
   9.750   1.0975   0.04569   0.04085  -0.0165   0.0094   1.0000
  10.000   1.0957   0.04985   0.04545  -0.0162   0.0092   1.0000
  10.250   1.0858   0.05444   0.05040  -0.0165   0.0091   1.0000
  10.500   1.0691   0.05945   0.05566  -0.0184   0.0091   1.0000
  10.750   1.0517   0.06677   0.06318  -0.0252   0.0092   1.0000
  11.000   1.0346   0.07635   0.07289  -0.0341   0.0094   1.0000
  11.250   0.9117   0.07222   0.06889  -0.0238   0.0101   1.0000
  11.500   0.8707   0.08636   0.08316  -0.0309   0.0104   1.0000
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