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AG14 (ag14-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: AG14 (ag14-il)
Reynolds number: 200,000
Max Cl/Cd: 59.38 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag14-il-200000-n5.txt
Download as CSV file: xf-ag14-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG14                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5595   0.11512   0.11157   0.0184   1.0000   0.0163
  -8.750  -0.5541   0.11129   0.10776   0.0167   1.0000   0.0163
  -8.500  -0.5487   0.10745   0.10394   0.0149   1.0000   0.0163
  -8.250  -0.5432   0.10359   0.10011   0.0130   1.0000   0.0163
  -8.000  -0.5378   0.09971   0.09627   0.0109   1.0000   0.0163
  -7.750  -0.5322   0.09580   0.09238   0.0087   1.0000   0.0163
  -6.750  -0.5051   0.07812   0.07479  -0.0011   1.0000   0.0140
  -6.500  -0.4910   0.07289   0.06957  -0.0068   1.0000   0.0128
  -6.250  -0.4725   0.06657   0.06322  -0.0143   1.0000   0.0117
  -6.000  -0.4462   0.05705   0.05359  -0.0254   1.0000   0.0105
  -5.750  -0.4180   0.04897   0.04532  -0.0332   1.0000   0.0102
  -5.500  -0.3968   0.04643   0.04269  -0.0357   1.0000   0.0114
  -5.250  -0.3702   0.04201   0.03808  -0.0393   1.0000   0.0131
  -5.000  -0.3409   0.03587   0.03158  -0.0431   1.0000   0.0138
  -4.750  -0.3119   0.02993   0.02514  -0.0457   1.0000   0.0132
  -4.500  -0.2827   0.02494   0.01953  -0.0472   1.0000   0.0128
  -4.250  -0.2538   0.02123   0.01521  -0.0478   1.0000   0.0127
  -4.000  -0.2255   0.01858   0.01200  -0.0478   1.0000   0.0127
  -3.750  -0.1977   0.01663   0.00966  -0.0476   1.0000   0.0131
  -3.500  -0.1704   0.01510   0.00786  -0.0473   1.0000   0.0138
  -3.250  -0.1433   0.01389   0.00647  -0.0469   1.0000   0.0150
  -3.000  -0.1165   0.01290   0.00536  -0.0466   1.0000   0.0175
  -2.750  -0.0899   0.01219   0.00462  -0.0463   1.0000   0.0214
  -2.500  -0.0628   0.01153   0.00382  -0.0460   1.0000   0.0252
  -2.250  -0.0359   0.01094   0.00321  -0.0457   1.0000   0.0366
  -2.000  -0.0089   0.01041   0.00277  -0.0455   1.0000   0.0641
  -1.750   0.0177   0.00991   0.00253  -0.0455   1.0000   0.1231
  -1.500   0.0444   0.00947   0.00238  -0.0455   1.0000   0.2035
  -1.250   0.0721   0.00903   0.00227  -0.0457   0.9989   0.3053
  -1.000   0.1095   0.00849   0.00219  -0.0480   0.9858   0.4485
  -0.750   0.1438   0.00779   0.00214  -0.0494   0.9692   0.6367
  -0.500   0.1769   0.00697   0.00198  -0.0494   0.9461   1.0000
  -0.250   0.2135   0.00699   0.00185  -0.0509   0.9164   1.0000
   0.000   0.2464   0.00703   0.00173  -0.0515   0.8791   1.0000
   0.250   0.2749   0.00712   0.00164  -0.0511   0.8370   1.0000
   0.500   0.3009   0.00727   0.00159  -0.0502   0.7929   1.0000
   0.750   0.3263   0.00747   0.00155  -0.0492   0.7482   1.0000
   1.000   0.3519   0.00769   0.00155  -0.0483   0.7035   1.0000
   1.250   0.3776   0.00794   0.00157  -0.0476   0.6588   1.0000
   1.500   0.4035   0.00821   0.00161  -0.0470   0.6149   1.0000
   1.750   0.4296   0.00849   0.00170  -0.0464   0.5723   1.0000
   2.000   0.4559   0.00877   0.00179  -0.0460   0.5312   1.0000
   2.250   0.4823   0.00906   0.00190  -0.0456   0.4919   1.0000
   2.500   0.5089   0.00935   0.00203  -0.0453   0.4546   1.0000
   2.750   0.5354   0.00966   0.00218  -0.0450   0.4186   1.0000
   3.000   0.5620   0.00997   0.00240  -0.0447   0.3834   1.0000
   3.250   0.5885   0.01030   0.00260  -0.0445   0.3504   1.0000
   3.500   0.6151   0.01063   0.00282  -0.0443   0.3177   1.0000
   3.750   0.6416   0.01098   0.00307  -0.0440   0.2865   1.0000
   4.000   0.6680   0.01135   0.00335  -0.0438   0.2566   1.0000
   4.250   0.6944   0.01173   0.00369  -0.0436   0.2278   1.0000
   4.500   0.7208   0.01214   0.00403  -0.0434   0.2006   1.0000
   4.750   0.7470   0.01258   0.00441  -0.0432   0.1741   1.0000
   5.000   0.7731   0.01303   0.00482  -0.0430   0.1510   1.0000
   5.250   0.7991   0.01351   0.00527  -0.0427   0.1285   1.0000
   5.500   0.8249   0.01405   0.00581  -0.0425   0.1097   1.0000
   5.750   0.8507   0.01458   0.00635  -0.0422   0.0924   1.0000
   6.000   0.8762   0.01515   0.00694  -0.0419   0.0771   1.0000
   6.250   0.9016   0.01577   0.00759  -0.0416   0.0637   1.0000
   6.500   0.9267   0.01644   0.00830  -0.0412   0.0513   1.0000
   6.750   0.9515   0.01718   0.00910  -0.0409   0.0405   1.0000
   7.000   0.9758   0.01804   0.01007  -0.0404   0.0313   1.0000
   7.250   0.9994   0.01905   0.01113  -0.0400   0.0233   1.0000
   7.500   1.0231   0.02005   0.01229  -0.0394   0.0184   1.0000
   7.750   1.0442   0.02167   0.01406  -0.0386   0.0150   1.0000
   8.000   1.0673   0.02269   0.01527  -0.0380   0.0123   1.0000
   8.250   1.0889   0.02399   0.01674  -0.0374   0.0106   1.0000
   8.500   1.1061   0.02636   0.01935  -0.0364   0.0094   1.0000
   8.750   1.1240   0.02855   0.02184  -0.0353   0.0088   1.0000
   9.000   1.1409   0.03092   0.02461  -0.0342   0.0084   1.0000
   9.250   1.1553   0.03372   0.02781  -0.0331   0.0080   1.0000
   9.500   1.1663   0.03698   0.03150  -0.0319   0.0077   1.0000
   9.750   1.1729   0.04072   0.03569  -0.0307   0.0075   1.0000
  10.000   1.1746   0.04470   0.04012  -0.0297   0.0074   1.0000
  10.250   1.1718   0.04871   0.04450  -0.0289   0.0071   1.0000
  10.500   1.1620   0.05278   0.04889  -0.0283   0.0070   1.0000
  10.750   1.1453   0.05763   0.05400  -0.0296   0.0070   1.0000
  11.000   1.1253   0.06497   0.06159  -0.0354   0.0071   1.0000
  11.250   1.1028   0.07555   0.07238  -0.0450   0.0074   1.0000
  11.500   1.0799   0.08727   0.08423  -0.0541   0.0077   1.0000
  11.750   1.0559   0.09896   0.09599  -0.0615   0.0080   1.0000
  12.000   1.0301   0.11083   0.10781  -0.0678   0.0082   1.0000
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