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NASA SC(2)-0606 AIRFOIL (sc20606-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0606 AIRFOIL (sc20606-il)
Reynolds number: 100,000
Max Cl/Cd: 35.04 at α=1.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20606-il-100000-n5.txt
Download as CSV file: xf-sc20606-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0606 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.5005   0.09589   0.09086  -0.0135   1.0000   0.0314
 -10.000  -0.5048   0.09073   0.08574  -0.0152   1.0000   0.0299
  -9.000  -0.6189   0.07812   0.07306  -0.0239   1.0000   0.0255
  -8.750  -0.6195   0.07242   0.06734  -0.0289   1.0000   0.0250
  -8.500  -0.6186   0.06689   0.06174  -0.0330   1.0000   0.0247
  -8.250  -0.6151   0.06140   0.05610  -0.0365   1.0000   0.0243
  -8.000  -0.6079   0.05593   0.05042  -0.0397   1.0000   0.0240
  -7.750  -0.5964   0.05050   0.04468  -0.0425   1.0000   0.0237
  -7.500  -0.5803   0.04519   0.03898  -0.0450   1.0000   0.0236
  -7.250  -0.5595   0.04016   0.03345  -0.0473   1.0000   0.0236
  -7.000  -0.5343   0.03588   0.02853  -0.0492   1.0000   0.0247
  -6.750  -0.5071   0.03225   0.02424  -0.0508   1.0000   0.0264
  -6.500  -0.4802   0.02883   0.02037  -0.0520   1.0000   0.0272
  -6.250  -0.4527   0.02617   0.01726  -0.0527   1.0000   0.0278
  -6.000  -0.4254   0.02403   0.01481  -0.0529   1.0000   0.0287
  -5.750  -0.3983   0.02223   0.01278  -0.0530   1.0000   0.0300
  -5.500  -0.3714   0.02068   0.01104  -0.0529   1.0000   0.0319
  -5.250  -0.3444   0.01944   0.00963  -0.0527   1.0000   0.0354
  -5.000  -0.3171   0.01825   0.00836  -0.0530   1.0000   0.0412
  -4.750  -0.2887   0.01721   0.00724  -0.0533   1.0000   0.0477
  -4.500  -0.2597   0.01628   0.00628  -0.0539   1.0000   0.0612
  -4.250  -0.2297   0.01535   0.00542  -0.0548   1.0000   0.0850
  -4.000  -0.1981   0.01421   0.00470  -0.0564   1.0000   0.1647
  -3.750  -0.1654   0.01243   0.00439  -0.0588   1.0000   0.4971
  -3.500  -0.1420   0.01223   0.00453  -0.0574   1.0000   0.6199
  -3.250  -0.1190   0.01225   0.00463  -0.0558   1.0000   0.6834
  -3.000  -0.0989   0.01236   0.00480  -0.0534   1.0000   0.7299
  -2.750  -0.0788   0.01247   0.00491  -0.0510   1.0000   0.7638
  -2.500  -0.0573   0.01250   0.00489  -0.0492   1.0000   0.7863
  -2.250  -0.0356   0.01251   0.00487  -0.0475   1.0000   0.8059
  -2.000  -0.0127   0.01249   0.00482  -0.0463   1.0000   0.8219
  -1.750   0.0109   0.01245   0.00475  -0.0453   1.0000   0.8346
  -1.500   0.0349   0.01240   0.00468  -0.0444   1.0000   0.8456
  -1.250   0.0594   0.01235   0.00463  -0.0438   1.0000   0.8567
  -1.000   0.0843   0.01230   0.00459  -0.0432   1.0000   0.8677
  -0.750   0.1095   0.01226   0.00458  -0.0428   1.0000   0.8787
  -0.500   0.1340   0.01222   0.00459  -0.0422   1.0000   0.8898
  -0.250   0.1577   0.01217   0.00461  -0.0414   1.0000   0.9018
   0.000   0.1815   0.01211   0.00465  -0.0407   1.0000   0.9152
   0.250   0.2049   0.01205   0.00469  -0.0399   1.0000   0.9304
   0.500   0.2280   0.01198   0.00475  -0.0391   1.0000   0.9492
   0.750   0.2467   0.01192   0.00485  -0.0378   1.0000   1.0000
   1.000   0.2781   0.01212   0.00522  -0.0391   1.0000   1.0000
   1.250   0.3331   0.01207   0.00540  -0.0448   0.9770   1.0000
   1.500   0.3967   0.01132   0.00489  -0.0504   0.8815   1.0000
   1.750   0.4512   0.01339   0.00458  -0.0537   0.2864   1.0000
   2.000   0.4769   0.01506   0.00531  -0.0541   0.1064   1.0000
   2.250   0.5051   0.01612   0.00617  -0.0545   0.0686   1.0000
   2.500   0.5330   0.01711   0.00710  -0.0548   0.0503   1.0000
   2.750   0.5608   0.01823   0.00834  -0.0547   0.0430   1.0000
   3.000   0.5875   0.01968   0.00977  -0.0548   0.0368   1.0000
   3.250   0.6164   0.02072   0.01100  -0.0549   0.0322   1.0000
   3.500   0.6450   0.02231   0.01274  -0.0549   0.0300   1.0000
   3.750   0.6736   0.02410   0.01473  -0.0549   0.0285   1.0000
   4.000   0.7018   0.02616   0.01704  -0.0548   0.0274   1.0000
   4.250   0.7289   0.02859   0.01986  -0.0545   0.0267   1.0000
   4.500   0.7545   0.03148   0.02318  -0.0541   0.0263   1.0000
   4.750   0.7768   0.03514   0.02732  -0.0535   0.0250   1.0000
   5.000   0.8005   0.03800   0.03086  -0.0524   0.0233   1.0000
   5.250   0.8202   0.04208   0.03550  -0.0514   0.0229   1.0000
   5.500   0.8372   0.04657   0.04049  -0.0504   0.0229   1.0000
   5.750   0.8516   0.05138   0.04574  -0.0496   0.0231   1.0000
   6.000   0.8635   0.05639   0.05112  -0.0490   0.0233   1.0000
   6.250   0.8727   0.06156   0.05661  -0.0487   0.0237   1.0000
   6.500   0.8790   0.06681   0.06211  -0.0487   0.0241   1.0000
   6.750   0.8824   0.07209   0.06759  -0.0489   0.0245   1.0000
   7.000   0.8824   0.07743   0.07302  -0.0492   0.0251   1.0000
   7.250   0.8795   0.08742   0.08343  -0.0548   0.0300   1.0000
   7.500   0.8726   0.09361   0.08968  -0.0581   0.0317   1.0000
   7.750   0.8662   0.09944   0.09551  -0.0619   0.0333   1.0000
   8.250   0.8703   0.11147   0.10754  -0.0600   0.0490   1.0000
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