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NASA SC(2)-1006 AIRFOIL (sc21006-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-1006 AIRFOIL (sc21006-il)
Reynolds number: 50,000
Max Cl/Cd: 29.81 at α=0.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc21006-il-50000-n5.txt
Download as CSV file: xf-sc21006-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-1006 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.750  -0.5167   0.11436   0.10703  -0.0210   0.9999   0.1164
 -11.500  -0.5305   0.11260   0.10542  -0.0233   0.9999   0.1177
 -11.250  -0.5272   0.10851   0.10140  -0.0230   0.9999   0.1186
 -11.000  -0.5138   0.10010   0.09289  -0.0246   0.9999   0.0679
 -10.750  -0.5163   0.09608   0.08893  -0.0251   0.9999   0.0590
 -10.500  -0.5205   0.09256   0.08549  -0.0253   0.9999   0.0565
 -10.000  -0.5249   0.08133   0.07449  -0.0361   0.9999   0.0497
  -9.750  -0.5198   0.07691   0.07009  -0.0383   0.9999   0.0488
  -9.500  -0.5130   0.07178   0.06487  -0.0421   0.9999   0.0478
  -9.250  -0.5032   0.06618   0.05920  -0.0470   0.9999   0.0469
  -9.000  -0.4885   0.06009   0.05292  -0.0529   0.9999   0.0459
  -8.750  -0.4668   0.05360   0.04609  -0.0598   0.9999   0.0450
  -8.500  -0.4372   0.04719   0.03914  -0.0669   0.9999   0.0444
  -8.250  -0.4026   0.04156   0.03285  -0.0730   0.9999   0.0444
  -8.000  -0.3662   0.03704   0.02762  -0.0777   0.9999   0.0460
  -7.750  -0.3287   0.03353   0.02321  -0.0813   0.9999   0.0500
  -7.500  -0.2968   0.03062   0.01993  -0.0833   0.9999   0.0528
  -7.250  -0.2663   0.02840   0.01737  -0.0842   0.9999   0.0555
  -7.000  -0.2369   0.02654   0.01512  -0.0845   0.9999   0.0594
  -6.750  -0.2084   0.02512   0.01353  -0.0849   0.9999   0.0688
  -6.500  -0.1801   0.02380   0.01201  -0.0849   0.9999   0.0788
  -6.250  -0.1482   0.02245   0.01057  -0.0860   0.9999   0.0941
  -6.000  -0.1090   0.02079   0.00919  -0.0894   0.9999   0.1546
  -5.750  -0.0668   0.01890   0.00868  -0.0939   0.9999   0.4385
  -5.500  -0.0447   0.01871   0.00888  -0.0919   0.9999   0.5734
  -5.250  -0.0254   0.01873   0.00895  -0.0893   0.9999   0.6429
  -5.000  -0.0073   0.01877   0.00896  -0.0864   0.9999   0.6952
  -4.750   0.0112   0.01879   0.00890  -0.0837   0.9999   0.7418
  -4.500   0.0289   0.01871   0.00876  -0.0810   0.9999   0.7804
  -4.250   0.0476   0.01856   0.00848  -0.0786   0.9999   0.8146
  -4.000   0.0672   0.01835   0.00817  -0.0767   0.9999   0.8478
  -3.750   0.0860   0.01804   0.00780  -0.0747   0.9999   0.8814
  -3.500   0.1041   0.01763   0.00736  -0.0727   0.9999   0.9203
  -3.250   0.1175   0.01722   0.00697  -0.0702   0.9999   1.0001
  -3.000   0.1530   0.01741   0.00699  -0.0727   0.9999   1.0001
  -2.750   0.1870   0.01762   0.00707  -0.0748   0.9999   1.0001
  -2.500   0.2197   0.01784   0.00722  -0.0765   0.9999   1.0001
  -2.250   0.2515   0.01808   0.00743  -0.0780   0.9999   1.0001
  -2.000   0.2824   0.01835   0.00769  -0.0793   0.9999   1.0001
  -1.750   0.3126   0.01863   0.00799  -0.0804   0.9999   1.0001
  -1.500   0.3421   0.01894   0.00835  -0.0813   0.9999   1.0001
  -1.250   0.3710   0.01926   0.00877  -0.0821   0.9999   1.0001
  -1.000   0.3994   0.01961   0.00928  -0.0828   0.9999   1.0001
  -0.750   0.4273   0.01999   0.00982  -0.0834   0.9999   1.0001
  -0.500   0.4548   0.02040   0.01041  -0.0839   0.9999   1.0001
  -0.250   0.4818   0.02084   0.01108  -0.0843   0.9999   1.0001
   0.000   0.5084   0.02132   0.01190  -0.0846   0.9999   1.0001
   0.250   0.5346   0.02185   0.01274  -0.0849   0.9999   1.0001
   0.500   0.5603   0.02242   0.01370  -0.0852   0.9999   1.0001
   0.750   0.6982   0.02342   0.01115  -0.0972   0.0969   1.0001
   1.000   0.7261   0.02496   0.01270  -0.0970   0.0756   1.0001
   1.250   0.7560   0.02655   0.01440  -0.0971   0.0642   1.0001
   1.500   0.7891   0.02831   0.01638  -0.0973   0.0584   1.0001
   1.750   0.8213   0.03043   0.01857  -0.0976   0.0524   1.0001
   2.000   0.8525   0.03240   0.02087  -0.0976   0.0469   1.0001
   2.250   0.8831   0.03496   0.02381  -0.0974   0.0449   1.0001
   2.500   0.9113   0.03781   0.02703  -0.0970   0.0435   1.0001
   2.750   0.9372   0.04102   0.03065  -0.0962   0.0427   1.0001
   3.000   0.9603   0.04455   0.03465  -0.0952   0.0422   1.0001
   3.250   0.9791   0.04871   0.03921  -0.0939   0.0411   1.0001
   3.500   0.9966   0.05197   0.04323  -0.0919   0.0396   1.0001
   3.750   1.0111   0.05587   0.04775  -0.0899   0.0384   1.0001
   4.000   1.0224   0.06020   0.05258  -0.0881   0.0382   1.0001
   4.250   1.0310   0.06475   0.05755  -0.0863   0.0386   1.0001
   4.500   1.0368   0.06944   0.06261  -0.0846   0.0390   1.0001
   4.750   1.0400   0.07421   0.06769  -0.0832   0.0395   1.0001
   5.000   1.0412   0.07905   0.07286  -0.0819   0.0400   1.0001
   5.250   1.0415   0.08396   0.07793  -0.0808   0.0405   1.0001
   5.750   0.9271   0.08520   0.07965  -0.0629   0.0410   1.0001
   6.250   0.9884   0.10616   0.10084  -0.0834   0.0467   1.0001
   6.500   0.9782   0.11279   0.10747  -0.0876   0.0482   1.0001
   6.750   0.9725   0.11950   0.11416  -0.0923   0.0498   1.0001
   7.000   0.9711   0.12525   0.11990  -0.0947   0.0518   1.0001
   7.500   0.7987   0.12743   0.12223  -0.0799   0.0497   1.0001
   7.750   0.7948   0.13203   0.12682  -0.0811   0.0514   1.0001
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