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AG17 (ag17-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: AG17 (ag17-il)
Reynolds number: 1,000,000
Max Cl/Cd: 95.42 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag17-il-1000000.txt
Download as CSV file: xf-ag17-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG17                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5563   0.09382   0.09229   0.0125   1.0000   0.0077
  -8.000  -0.5529   0.08972   0.08819   0.0101   1.0000   0.0079
  -7.750  -0.5495   0.08551   0.08400   0.0073   1.0000   0.0081
  -7.500  -0.5447   0.08097   0.07949   0.0033   1.0000   0.0084
  -7.250  -0.5334   0.07549   0.07401  -0.0031   1.0000   0.0087
  -7.000  -0.5168   0.06983   0.06835  -0.0103   1.0000   0.0092
  -6.750  -0.4959   0.06406   0.06256  -0.0180   1.0000   0.0095
  -5.750  -0.4065   0.01880   0.01525  -0.0505   1.0000   0.0059
  -5.500  -0.3803   0.01718   0.01338  -0.0505   1.0000   0.0057
  -5.250  -0.3545   0.01292   0.00849  -0.0507   1.0000   0.0060
  -5.000  -0.3284   0.01145   0.00686  -0.0505   1.0000   0.0068
  -4.750  -0.3021   0.01090   0.00624  -0.0502   1.0000   0.0074
  -4.500  -0.2758   0.01022   0.00548  -0.0498   1.0000   0.0079
  -4.250  -0.2496   0.00953   0.00470  -0.0494   1.0000   0.0083
  -4.000  -0.2234   0.00900   0.00410  -0.0490   1.0000   0.0088
  -3.750  -0.1976   0.00873   0.00378  -0.0485   1.0000   0.0095
  -3.500  -0.1691   0.00816   0.00313  -0.0486   0.9994   0.0098
  -3.250  -0.1324   0.00728   0.00210  -0.0503   0.9962   0.0137
  -3.000  -0.0960   0.00687   0.00169  -0.0520   0.9921   0.0245
  -2.750  -0.0610   0.00650   0.00145  -0.0536   0.9849   0.0567
  -2.500  -0.0268   0.00617   0.00127  -0.0550   0.9748   0.1003
  -2.250   0.0056   0.00586   0.00112  -0.0560   0.9598   0.1528
  -2.000   0.0341   0.00561   0.00101  -0.0560   0.9390   0.2099
  -1.750   0.0600   0.00541   0.00092  -0.0553   0.9139   0.2676
  -1.500   0.0859   0.00530   0.00085  -0.0547   0.8870   0.3141
  -1.250   0.1121   0.00521   0.00079  -0.0542   0.8588   0.3637
  -1.000   0.1388   0.00514   0.00075  -0.0538   0.8300   0.4135
  -0.750   0.1657   0.00510   0.00072  -0.0534   0.8009   0.4588
  -0.500   0.1928   0.00504   0.00070  -0.0532   0.7710   0.5166
  -0.250   0.2199   0.00496   0.00070  -0.0529   0.7414   0.5812
   0.000   0.2469   0.00486   0.00071  -0.0527   0.7120   0.6570
   0.250   0.2729   0.00463   0.00074  -0.0522   0.6828   0.7752
   0.750   0.3252   0.00440   0.00073  -0.0509   0.6217   1.0000
   1.000   0.3529   0.00456   0.00075  -0.0508   0.5895   1.0000
   1.250   0.3806   0.00473   0.00079  -0.0506   0.5572   1.0000
   1.500   0.4083   0.00490   0.00083  -0.0506   0.5242   1.0000
   1.750   0.4360   0.00510   0.00089  -0.0505   0.4908   1.0000
   2.000   0.4637   0.00529   0.00095  -0.0504   0.4583   1.0000
   2.250   0.4913   0.00550   0.00103  -0.0503   0.4246   1.0000
   2.500   0.5188   0.00572   0.00112  -0.0502   0.3926   1.0000
   2.750   0.5464   0.00594   0.00123  -0.0501   0.3609   1.0000
   3.000   0.5738   0.00618   0.00135  -0.0501   0.3293   1.0000
   3.250   0.6013   0.00641   0.00147  -0.0500   0.3014   1.0000
   3.500   0.6287   0.00666   0.00161  -0.0499   0.2721   1.0000
   3.750   0.6560   0.00691   0.00176  -0.0498   0.2458   1.0000
   4.000   0.6832   0.00718   0.00194  -0.0497   0.2183   1.0000
   4.250   0.7105   0.00745   0.00212  -0.0496   0.1943   1.0000
   4.500   0.7376   0.00773   0.00232  -0.0495   0.1717   1.0000
   4.750   0.7645   0.00806   0.00254  -0.0494   0.1465   1.0000
   5.000   0.7914   0.00839   0.00279  -0.0493   0.1235   1.0000
   5.250   0.8182   0.00871   0.00305  -0.0491   0.1039   1.0000
   5.500   0.8449   0.00907   0.00333  -0.0490   0.0851   1.0000
   5.750   0.8714   0.00946   0.00364  -0.0488   0.0672   1.0000
   6.000   0.8978   0.00986   0.00397  -0.0486   0.0513   1.0000
   6.250   0.9240   0.01029   0.00433  -0.0484   0.0381   1.0000
   6.500   0.9502   0.01070   0.00473  -0.0482   0.0287   1.0000
   6.750   0.9763   0.01114   0.00516  -0.0480   0.0210   1.0000
   7.000   1.0020   0.01169   0.00571  -0.0477   0.0140   1.0000
   7.250   1.0268   0.01250   0.00655  -0.0472   0.0085   1.0000
   7.500   1.0527   0.01291   0.00700  -0.0469   0.0070   1.0000
   7.750   1.0770   0.01376   0.00792  -0.0464   0.0052   1.0000
   8.000   1.1006   0.01477   0.00909  -0.0457   0.0047   1.0000
   8.250   1.1250   0.01547   0.00992  -0.0453   0.0045   1.0000
   8.500   1.1488   0.01628   0.01084  -0.0447   0.0042   1.0000
   8.750   1.1721   0.01715   0.01183  -0.0441   0.0040   1.0000
   9.000   1.1951   0.01805   0.01284  -0.0435   0.0037   1.0000
   9.250   1.2178   0.01892   0.01381  -0.0430   0.0034   1.0000
   9.500   1.2399   0.01986   0.01485  -0.0424   0.0032   1.0000
   9.750   1.2564   0.02187   0.01706  -0.0412   0.0028   1.0000
  10.000   1.2597   0.02640   0.02211  -0.0387   0.0025   1.0000
  10.250   1.2764   0.02801   0.02392  -0.0377   0.0025   1.0000
  10.500   1.2884   0.03027   0.02643  -0.0363   0.0024   1.0000
  10.750   1.2954   0.03306   0.02950  -0.0347   0.0024   1.0000
  11.000   1.2965   0.03629   0.03305  -0.0329   0.0024   1.0000
  11.250   1.2906   0.03978   0.03682  -0.0309   0.0024   1.0000
  11.500   1.2752   0.04334   0.04061  -0.0288   0.0024   1.0000
  11.750   1.2578   0.04806   0.04554  -0.0296   0.0024   1.0000
  12.000   1.2404   0.05442   0.05212  -0.0338   0.0024   1.0000
  12.250   1.2220   0.06277   0.06066  -0.0406   0.0024   1.0000
  12.500   1.2018   0.07262   0.07065  -0.0483   0.0025   1.0000
  12.750   1.1780   0.08341   0.08158  -0.0556   0.0025   1.0000
  13.000   1.1530   0.09421   0.09249  -0.0619   0.0025   1.0000
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