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S6062 8% (s6062-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: S6062 8% (s6062-il)
Reynolds number: 500,000
Max Cl/Cd: 81.99 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s6062-il-500000.txt
Download as CSV file: xf-s6062-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.4512   0.09244   0.09037  -0.0081   1.0000   0.0163
  -9.500  -0.4485   0.08874   0.08669  -0.0090   1.0000   0.0167
  -9.250  -0.4443   0.08544   0.08339  -0.0098   1.0000   0.0172
  -9.000  -0.4427   0.08153   0.07949  -0.0110   1.0000   0.0178
  -8.750  -0.4421   0.07755   0.07552  -0.0125   1.0000   0.0187
  -8.500  -0.4458   0.07242   0.07041  -0.0146   1.0000   0.0185
  -8.250  -0.4482   0.06794   0.06595  -0.0167   1.0000   0.0197
  -8.000  -0.4539   0.06286   0.06089  -0.0193   1.0000   0.0198
  -7.750  -0.4629   0.05762   0.05569  -0.0226   1.0000   0.0201
  -7.500  -0.4795   0.05134   0.04942  -0.0287   1.0000   0.0192
  -7.250  -0.4850   0.04480   0.04279  -0.0334   1.0000   0.0194
  -7.000  -0.4848   0.03873   0.03660  -0.0364   1.0000   0.0201
  -6.750  -0.4792   0.03295   0.03059  -0.0383   1.0000   0.0216
  -6.500  -0.4636   0.03050   0.02780  -0.0384   1.0000   0.0229
  -6.250  -0.4578   0.02710   0.02415  -0.0373   1.0000   0.0230
  -6.000  -0.4902   0.03271   0.02911  -0.0362   1.0000   0.0242
  -5.750  -0.4817   0.03068   0.02704  -0.0340   1.0000   0.0250
  -5.500  -0.4528   0.02845   0.02467  -0.0358   0.9970   0.0264
  -5.250  -0.4192   0.02606   0.02200  -0.0380   0.9934   0.0292
  -4.750  -0.3505   0.01593   0.01044  -0.0389   0.9858   0.0143
  -4.500  -0.3148   0.01363   0.00786  -0.0403   0.9835   0.0132
  -4.250  -0.2828   0.01225   0.00633  -0.0410   0.9778   0.0132
  -4.000  -0.2487   0.01125   0.00523  -0.0423   0.9730   0.0140
  -3.750  -0.2168   0.01055   0.00442  -0.0431   0.9666   0.0148
  -3.500  -0.1865   0.00998   0.00378  -0.0435   0.9589   0.0161
  -3.250  -0.1593   0.00942   0.00314  -0.0432   0.9495   0.0188
  -3.000  -0.1325   0.00909   0.00273  -0.0427   0.9405   0.0221
  -2.750  -0.1088   0.00828   0.00229  -0.0418   0.9302   0.1132
  -2.500  -0.0850   0.00772   0.00208  -0.0411   0.9194   0.2145
  -2.250  -0.0616   0.00709   0.00193  -0.0405   0.9092   0.3509
  -2.000  -0.0385   0.00649   0.00181  -0.0396   0.8992   0.4980
  -1.750  -0.0160   0.00595   0.00174  -0.0385   0.8885   0.6427
  -1.500   0.0070   0.00562   0.00172  -0.0372   0.8774   0.7441
  -1.250   0.0300   0.00544   0.00171  -0.0356   0.8666   0.8170
  -1.000   0.0539   0.00536   0.00168  -0.0343   0.8558   0.8589
  -0.750   0.0776   0.00531   0.00165  -0.0329   0.8448   0.8935
  -0.500   0.1022   0.00528   0.00161  -0.0317   0.8327   0.9244
  -0.250   0.1309   0.00528   0.00157  -0.0315   0.8207   0.9495
   0.000   0.1648   0.00529   0.00152  -0.0325   0.8084   0.9687
   0.250   0.2029   0.00531   0.00147  -0.0346   0.7954   0.9821
   0.500   0.2445   0.00532   0.00141  -0.0374   0.7809   0.9912
   0.750   0.2857   0.00534   0.00135  -0.0403   0.7653   1.0000
   1.000   0.3091   0.00539   0.00132  -0.0394   0.7491   1.0000
   1.500   0.3575   0.00552   0.00132  -0.0379   0.7147   1.0000
   1.750   0.3822   0.00561   0.00133  -0.0372   0.6959   1.0000
   2.250   0.4325   0.00582   0.00141  -0.0361   0.6552   1.0000
   2.500   0.4580   0.00596   0.00148  -0.0355   0.6327   1.0000
   2.750   0.4836   0.00611   0.00155  -0.0351   0.6077   1.0000
   3.000   0.5095   0.00627   0.00163  -0.0347   0.5814   1.0000
   3.250   0.5346   0.00652   0.00173  -0.0341   0.5389   1.0000
   3.500   0.5593   0.00685   0.00186  -0.0336   0.4849   1.0000
   3.750   0.5835   0.00729   0.00203  -0.0331   0.4196   1.0000
   4.000   0.6083   0.00770   0.00223  -0.0327   0.3648   1.0000
   4.250   0.6330   0.00817   0.00247  -0.0323   0.3078   1.0000
   4.500   0.6568   0.00879   0.00279  -0.0319   0.2361   1.0000
   4.750   0.6794   0.00962   0.00319  -0.0314   0.1496   1.0000
   5.000   0.7014   0.01061   0.00373  -0.0308   0.0627   1.0000
   5.250   0.7246   0.01145   0.00435  -0.0302   0.0230   1.0000
   5.500   0.7499   0.01191   0.00487  -0.0298   0.0201   1.0000
   5.750   0.7746   0.01252   0.00555  -0.0292   0.0179   1.0000
   6.000   0.7982   0.01327   0.00642  -0.0285   0.0171   1.0000
   6.250   0.8200   0.01431   0.00758  -0.0275   0.0164   1.0000
   6.500   0.8440   0.01492   0.00826  -0.0270   0.0159   1.0000
   6.750   0.8663   0.01583   0.00927  -0.0261   0.0156   1.0000
   7.000   0.8883   0.01686   0.01042  -0.0252   0.0153   1.0000
   7.250   0.9099   0.01803   0.01170  -0.0243   0.0150   1.0000
   7.500   0.9314   0.01941   0.01320  -0.0233   0.0148   1.0000
   7.750   0.9527   0.02111   0.01505  -0.0222   0.0147   1.0000
   8.000   0.9736   0.02327   0.01741  -0.0211   0.0149   1.0000
   8.250   0.9930   0.02601   0.02043  -0.0199   0.0153   1.0000
   8.500   1.0115   0.02838   0.02307  -0.0187   0.0152   1.0000
   8.750   1.0253   0.03180   0.02686  -0.0171   0.0155   1.0000
  10.250   0.9998   0.05842   0.05570  -0.0027   0.0210   1.0000
  10.500   0.9797   0.06205   0.05948  -0.0016   0.0209   1.0000
  10.750   0.9579   0.06664   0.06422  -0.0027   0.0208   1.0000
  11.000   0.9342   0.07260   0.07030  -0.0059   0.0214   1.0000
  11.250   0.9086   0.08038   0.07820  -0.0117   0.0220   1.0000
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