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S6062 8% (s6062-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: S6062 8% (s6062-il)
Reynolds number: 200,000
Max Cl/Cd: 58.82 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-s6062-il-200000-n5.txt
Download as CSV file: xf-s6062-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: S6062 8%                                        
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5262   0.08314   0.07973  -0.0144   1.0000   0.0164
  -8.250  -0.5279   0.07883   0.07546  -0.0169   1.0000   0.0157
  -8.000  -0.5323   0.07413   0.07083  -0.0204   1.0000   0.0152
  -7.750  -0.5367   0.06815   0.06486  -0.0264   1.0000   0.0146
  -7.500  -0.5381   0.06116   0.05779  -0.0318   1.0000   0.0139
  -7.000  -0.5323   0.04482   0.04083  -0.0377   1.0000   0.0113
  -6.750  -0.5250   0.04016   0.03587  -0.0380   1.0000   0.0111
  -6.500  -0.5148   0.03589   0.03125  -0.0378   1.0000   0.0109
  -6.250  -0.5025   0.03203   0.02699  -0.0369   1.0000   0.0107
  -6.000  -0.4888   0.02873   0.02326  -0.0356   1.0000   0.0106
  -5.750  -0.4737   0.02593   0.02004  -0.0340   1.0000   0.0105
  -5.500  -0.4573   0.02353   0.01724  -0.0323   1.0000   0.0105
  -5.250  -0.4362   0.02146   0.01479  -0.0313   0.9986   0.0105
  -5.000  -0.4032   0.01942   0.01236  -0.0326   0.9930   0.0106
  -4.750  -0.3699   0.01777   0.01040  -0.0338   0.9875   0.0110
  -4.500  -0.3351   0.01659   0.00903  -0.0352   0.9829   0.0116
  -4.250  -0.3044   0.01517   0.00753  -0.0363   0.9761   0.0136
  -3.750  -0.2392   0.01354   0.00567  -0.0384   0.9626   0.0156
  -3.500  -0.2053   0.01294   0.00488  -0.0397   0.9566   0.0171
  -3.250  -0.1750   0.01238   0.00427  -0.0401   0.9477   0.0208
  -3.000  -0.1445   0.01169   0.00374  -0.0408   0.9396   0.0593
  -2.750  -0.1161   0.01091   0.00342  -0.0413   0.9308   0.1614
  -2.500  -0.0893   0.01033   0.00316  -0.0414   0.9205   0.2573
  -2.250  -0.0650   0.00948   0.00301  -0.0411   0.9104   0.4322
  -2.000  -0.0412   0.00889   0.00290  -0.0402   0.9005   0.5686
  -1.750  -0.0180   0.00853   0.00285  -0.0388   0.8896   0.6688
  -1.500   0.0050   0.00830   0.00281  -0.0372   0.8785   0.7489
  -1.250   0.0286   0.00817   0.00277  -0.0356   0.8676   0.8102
  -1.000   0.0534   0.00810   0.00272  -0.0343   0.8569   0.8575
  -0.750   0.0810   0.00805   0.00264  -0.0336   0.8459   0.8949
  -0.500   0.1125   0.00802   0.00255  -0.0339   0.8341   0.9263
  -0.250   0.1487   0.00800   0.00246  -0.0353   0.8223   0.9519
   0.000   0.1864   0.00798   0.00236  -0.0372   0.8097   0.9723
   0.250   0.2259   0.00795   0.00225  -0.0395   0.7962   0.9892
   0.500   0.2603   0.00796   0.00217  -0.0409   0.7814   1.0000
   0.750   0.2838   0.00800   0.00213  -0.0400   0.7658   1.0000
   1.000   0.3075   0.00806   0.00211  -0.0390   0.7494   1.0000
   1.250   0.3317   0.00814   0.00211  -0.0382   0.7322   1.0000
   1.500   0.3562   0.00822   0.00212  -0.0375   0.7134   1.0000
   1.750   0.3810   0.00832   0.00215  -0.0367   0.6939   1.0000
   2.000   0.4060   0.00843   0.00219  -0.0361   0.6738   1.0000
   2.250   0.4312   0.00856   0.00227  -0.0354   0.6523   1.0000
   2.500   0.4564   0.00871   0.00235  -0.0348   0.6295   1.0000
   2.750   0.4817   0.00887   0.00244  -0.0342   0.6053   1.0000
   3.000   0.5069   0.00906   0.00256  -0.0337   0.5791   1.0000
   3.250   0.5322   0.00927   0.00272  -0.0331   0.5503   1.0000
   3.500   0.5573   0.00951   0.00288  -0.0326   0.5185   1.0000
   3.750   0.5811   0.00988   0.00304  -0.0318   0.4667   1.0000
   4.000   0.6039   0.01039   0.00325  -0.0310   0.3969   1.0000
   4.250   0.6268   0.01097   0.00356  -0.0304   0.3288   1.0000
   4.500   0.6485   0.01177   0.00394  -0.0297   0.2424   1.0000
   4.750   0.6713   0.01248   0.00436  -0.0292   0.1790   1.0000
   5.000   0.6948   0.01313   0.00482  -0.0287   0.1319   1.0000
   5.250   0.7165   0.01408   0.00543  -0.0282   0.0684   1.0000
   5.500   0.7386   0.01500   0.00614  -0.0275   0.0306   1.0000
   5.750   0.7621   0.01570   0.00687  -0.0269   0.0217   1.0000
   6.000   0.7856   0.01640   0.00764  -0.0263   0.0189   1.0000
   6.250   0.8090   0.01711   0.00846  -0.0256   0.0178   1.0000
   6.500   0.8319   0.01786   0.00937  -0.0249   0.0170   1.0000
   6.750   0.8539   0.01873   0.01038  -0.0241   0.0164   1.0000
   7.000   0.8752   0.01971   0.01148  -0.0233   0.0159   1.0000
   7.250   0.8958   0.02081   0.01275  -0.0223   0.0155   1.0000
   7.500   0.9160   0.02203   0.01408  -0.0213   0.0151   1.0000
   7.750   0.9361   0.02339   0.01557  -0.0203   0.0149   1.0000
   8.000   0.9563   0.02475   0.01707  -0.0193   0.0144   1.0000
   8.250   0.9761   0.02610   0.01853  -0.0185   0.0135   1.0000
   8.500   0.9946   0.02779   0.02036  -0.0176   0.0128   1.0000
   8.750   1.0124   0.02991   0.02269  -0.0166   0.0125   1.0000
   9.000   1.0275   0.03272   0.02581  -0.0154   0.0122   1.0000
   9.250   1.0387   0.03610   0.02958  -0.0139   0.0120   1.0000
   9.500   1.0493   0.03874   0.03258  -0.0124   0.0119   1.0000
   9.750   1.0562   0.04153   0.03574  -0.0108   0.0118   1.0000
  10.000   1.0584   0.04452   0.03908  -0.0089   0.0118   1.0000
  10.250   1.0554   0.04772   0.04262  -0.0069   0.0117   1.0000
  10.500   1.0454   0.05072   0.04589  -0.0044   0.0117   1.0000
  10.750   1.0319   0.05401   0.04942  -0.0026   0.0117   1.0000
  11.000   1.0164   0.05784   0.05347  -0.0021   0.0117   1.0000
  11.250   1.0004   0.06226   0.05808  -0.0030   0.0118   1.0000
  11.500   0.9824   0.06754   0.06356  -0.0053   0.0118   1.0000
  11.750   0.9640   0.07377   0.06996  -0.0091   0.0118   1.0000
  12.000   0.9451   0.08112   0.07745  -0.0142   0.0119   1.0000
  12.250   0.8367   0.08317   0.07981  -0.0140   0.0121   1.0000
  12.500   0.8172   0.09157   0.08835  -0.0190   0.0123   1.0000
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