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NASA SC(2)-0612 AIRFOIL (sc20612-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0612 AIRFOIL (sc20612-il)
Reynolds number: 100,000
Max Cl/Cd: 29.83 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20612-il-100000.txt
Download as CSV file: xf-sc20612-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0612 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.5741   0.10171   0.09609  -0.0334   1.0000   0.1811
 -10.250  -0.7918   0.06326   0.05684  -0.0540   1.0000   0.0903
 -10.000  -0.7915   0.05895   0.05243  -0.0533   1.0000   0.0894
  -9.750  -0.7952   0.05473   0.04794  -0.0528   1.0000   0.0891
  -9.500  -0.7938   0.05069   0.04356  -0.0524   1.0000   0.0892
  -9.250  -0.7869   0.04681   0.03927  -0.0523   1.0000   0.0895
  -9.000  -0.7747   0.04316   0.03518  -0.0524   1.0000   0.0898
  -8.750  -0.7580   0.03998   0.03149  -0.0525   1.0000   0.0905
  -8.500  -0.7400   0.03676   0.02809  -0.0526   1.0000   0.0927
  -8.250  -0.7197   0.03517   0.02655  -0.0521   1.0000   0.0960
  -8.000  -0.6976   0.03324   0.02435  -0.0520   1.0000   0.0994
  -7.750  -0.6736   0.03129   0.02198  -0.0521   1.0000   0.1025
  -7.500  -0.6508   0.02936   0.01993  -0.0519   1.0000   0.1068
  -7.250  -0.6276   0.02816   0.01870  -0.0515   1.0000   0.1122
  -7.000  -0.6022   0.02700   0.01725  -0.0513   1.0000   0.1175
  -6.750  -0.5796   0.02557   0.01596  -0.0507   1.0000   0.1238
  -6.500  -0.5546   0.02469   0.01494  -0.0504   1.0000   0.1313
  -6.250  -0.5311   0.02359   0.01397  -0.0499   1.0000   0.1390
  -6.000  -0.5055   0.02290   0.01317  -0.0497   1.0000   0.1483
  -5.750  -0.4815   0.02196   0.01244  -0.0492   1.0000   0.1585
  -5.500  -0.4560   0.02116   0.01176  -0.0490   1.0000   0.1708
  -5.250  -0.4295   0.02038   0.01116  -0.0491   1.0000   0.1879
  -5.000  -0.4008   0.01950   0.01059  -0.0498   1.0000   0.2170
  -4.750  -0.3572   0.01711   0.01021  -0.0547   1.0000   0.4683
  -4.500  -0.3480   0.01841   0.01190  -0.0487   1.0000   0.6002
  -4.250  -0.3311   0.01952   0.01293  -0.0453   1.0000   0.6393
  -4.000  -0.3205   0.02061   0.01399  -0.0402   1.0000   0.6586
  -3.750  -0.3081   0.02154   0.01488  -0.0358   1.0000   0.6757
  -3.500  -0.2948   0.02235   0.01564  -0.0317   1.0000   0.6920
  -3.250  -0.2806   0.02308   0.01633  -0.0280   1.0000   0.7086
  -3.000  -0.2742   0.02379   0.01705  -0.0222   1.0000   0.7206
  -2.750  -0.2642   0.02433   0.01758  -0.0175   1.0000   0.7349
  -2.500  -0.2512   0.02473   0.01795  -0.0138   1.0000   0.7503
  -2.250  -0.2390   0.02507   0.01828  -0.0100   1.0000   0.7665
  -2.000  -0.2289   0.02533   0.01854  -0.0056   1.0000   0.7834
  -1.750  -0.2203   0.02546   0.01866  -0.0010   1.0000   0.8006
  -1.500  -0.2112   0.02545   0.01866   0.0033   1.0000   0.8178
  -1.250  -0.1985   0.02534   0.01854   0.0064   1.0000   0.8335
  -1.000  -0.1757   0.02534   0.01850   0.0067   1.0000   0.8458
  -0.750  -0.1604   0.02499   0.01815   0.0087   1.0000   0.8533
  -0.500  -0.1313   0.02507   0.01820   0.0071   1.0000   0.8611
  -0.250  -0.1155   0.02476   0.01790   0.0090   1.0000   0.8687
   0.000  -0.0867   0.02490   0.01803   0.0074   1.0000   0.8760
   0.250  -0.0665   0.02472   0.01788   0.0079   1.0000   0.8809
   0.500  -0.0407   0.02480   0.01797   0.0069   1.0000   0.8852
   0.750  -0.0063   0.02507   0.01827   0.0041   0.9975   0.8891
   1.000   0.0438   0.02538   0.01862  -0.0012   0.9867   0.8927
   1.250   0.0906   0.02556   0.01885  -0.0056   0.9747   0.8961
   1.500   0.1534   0.02563   0.01899  -0.0126   0.9569   0.8984
   1.750   0.2118   0.02525   0.01870  -0.0180   0.9339   0.9004
   2.000   0.2756   0.02453   0.01808  -0.0238   0.9131   0.9025
   2.250   0.3311   0.02351   0.01719  -0.0275   0.8904   0.9049
   2.500   0.3845   0.02205   0.01588  -0.0301   0.8681   0.9073
   2.750   0.4372   0.02038   0.01437  -0.0320   0.8449   0.9097
   3.000   0.4777   0.01900   0.01312  -0.0321   0.8124   0.9126
   3.250   0.5155   0.01781   0.01199  -0.0316   0.7507   0.9151
   3.500   0.5477   0.01836   0.01076  -0.0294   0.4014   0.9168
   3.750   0.5592   0.02035   0.01152  -0.0273   0.2434   0.9191
   4.000   0.5793   0.02134   0.01212  -0.0262   0.2039   0.9216
   4.250   0.6033   0.02209   0.01272  -0.0257   0.1814   0.9246
   4.500   0.6300   0.02291   0.01342  -0.0258   0.1661   0.9280
   4.750   0.6594   0.02389   0.01427  -0.0265   0.1537   0.9311
   5.000   0.6892   0.02482   0.01509  -0.0273   0.1429   0.9339
   5.250   0.7170   0.02568   0.01605  -0.0274   0.1338   0.9374
   5.500   0.7475   0.02711   0.01729  -0.0283   0.1253   0.9409
   5.750   0.7767   0.02804   0.01849  -0.0286   0.1184   0.9442
   6.000   0.8067   0.02945   0.01982  -0.0296   0.1116   0.9470
   6.250   0.8334   0.03093   0.02161  -0.0294   0.1066   0.9505
   6.500   0.8600   0.03235   0.02326  -0.0294   0.1019   0.9546
   6.750   0.8877   0.03415   0.02502  -0.0301   0.0974   0.9588
   7.000   0.9095   0.03625   0.02757  -0.0294   0.0944   0.9632
   7.250   0.9309   0.03869   0.03050  -0.0286   0.0925   0.9677
   7.500   0.9505   0.04130   0.03355  -0.0279   0.0903   0.9725
   7.750   0.9704   0.04367   0.03622  -0.0275   0.0878   0.9782
   8.000   0.9939   0.04612   0.03879  -0.0279   0.0857   0.9866
   8.250   0.9966   0.05133   0.04483  -0.0258   0.0873   1.0000
   8.500   0.9891   0.05807   0.05231  -0.0235   0.0911   1.0000
   8.750   0.9892   0.06362   0.05821  -0.0227   0.0937   1.0000
   9.000   0.9980   0.06897   0.06368  -0.0230   0.0955   1.0000
   9.250   0.8266   0.10048   0.09657  -0.0332   0.1751   1.0000
   9.500   0.7707   0.11230   0.10832  -0.0471   0.1674   1.0000
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