Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NASA SC(2)-0606 AIRFOIL (sc20606-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0606 AIRFOIL (sc20606-il)
Reynolds number: 200,000
Max Cl/Cd: 51.78 at α=1.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20606-il-200000.txt
Download as CSV file: xf-sc20606-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.6052   0.09409   0.09045  -0.0125   1.0000   0.0426
  -9.250  -0.6095   0.08900   0.08542  -0.0185   1.0000   0.0432
  -9.000  -0.6134   0.08265   0.07908  -0.0272   1.0000   0.0434
  -8.750  -0.6140   0.07742   0.07373  -0.0334   1.0000   0.0436
  -8.500  -0.6126   0.07283   0.06895  -0.0372   1.0000   0.0438
  -8.250  -0.6149   0.06452   0.06070  -0.0392   1.0000   0.0449
  -8.000  -0.6051   0.06150   0.05776  -0.0381   1.0000   0.0464
  -7.750  -0.5941   0.05888   0.05513  -0.0382   1.0000   0.0496
  -7.500  -0.5739   0.05544   0.05094  -0.0445   1.0000   0.0565
  -7.250  -0.5637   0.04763   0.04311  -0.0468   1.0000   0.0585
  -7.000  -0.5461   0.04437   0.03987  -0.0472   1.0000   0.0608
  -6.750  -0.5186   0.04214   0.03709  -0.0493   1.0000   0.0691
  -6.500  -0.4972   0.03668   0.03149  -0.0516   1.0000   0.0732
  -6.250  -0.4568   0.02837   0.02209  -0.0541   1.0000   0.0465
  -6.000  -0.4228   0.02315   0.01616  -0.0550   1.0000   0.0357
  -5.750  -0.3936   0.02100   0.01368  -0.0554   1.0000   0.0370
  -5.500  -0.3650   0.01928   0.01169  -0.0556   1.0000   0.0389
  -5.250  -0.3362   0.01743   0.00958  -0.0555   1.0000   0.0395
  -5.000  -0.3078   0.01602   0.00803  -0.0554   1.0000   0.0415
  -4.750  -0.2786   0.01465   0.00657  -0.0556   1.0000   0.0449
  -4.500  -0.2491   0.01372   0.00564  -0.0563   1.0000   0.0554
  -4.250  -0.2157   0.01244   0.00443  -0.0579   1.0000   0.0736
  -4.000  -0.1793   0.01110   0.00354  -0.0606   1.0000   0.1598
  -3.750  -0.1388   0.00917   0.00349  -0.0647   1.0000   0.6206
  -3.500  -0.1136   0.00920   0.00357  -0.0637   1.0000   0.6823
  -3.250  -0.0887   0.00925   0.00364  -0.0627   1.0000   0.7143
  -3.000  -0.0653   0.00937   0.00377  -0.0613   1.0000   0.7435
  -2.750  -0.0422   0.00951   0.00391  -0.0598   1.0000   0.7693
  -2.500  -0.0218   0.00965   0.00408  -0.0575   1.0000   0.7917
  -2.250  -0.0005   0.00976   0.00421  -0.0556   1.0000   0.8116
  -2.000   0.0203   0.00983   0.00430  -0.0537   1.0000   0.8273
  -1.750   0.0417   0.00986   0.00435  -0.0520   1.0000   0.8419
  -1.500   0.0648   0.00986   0.00437  -0.0508   1.0000   0.8547
  -1.250   0.0892   0.00984   0.00437  -0.0500   1.0000   0.8666
  -1.000   0.1132   0.00981   0.00436  -0.0492   1.0000   0.8773
  -0.750   0.1364   0.00975   0.00436  -0.0481   1.0000   0.8876
  -0.500   0.1603   0.00970   0.00436  -0.0473   1.0000   0.8981
  -0.250   0.1846   0.00966   0.00438  -0.0467   1.0000   0.9091
   0.000   0.2083   0.00961   0.00440  -0.0460   1.0000   0.9209
   0.250   0.2308   0.00953   0.00441  -0.0449   1.0000   0.9340
   0.500   0.2516   0.00941   0.00440  -0.0436   1.0000   0.9499
   0.750   0.2712   0.00929   0.00440  -0.0421   1.0000   0.9749
   1.000   0.2969   0.00941   0.00466  -0.0424   1.0000   1.0000
   1.250   0.3643   0.00902   0.00448  -0.0504   0.9739   1.0000
   1.500   0.4246   0.00820   0.00387  -0.0556   0.9152   1.0000
   2.000   0.4852   0.01275   0.00468  -0.0561   0.0734   1.0000
   2.250   0.5148   0.01388   0.00572  -0.0567   0.0540   1.0000
   2.500   0.5446   0.01517   0.00704  -0.0571   0.0467   1.0000
   2.750   0.5735   0.01702   0.00881  -0.0576   0.0395   1.0000
   3.000   0.6049   0.01824   0.01023  -0.0581   0.0377   1.0000
   3.250   0.6361   0.02010   0.01231  -0.0583   0.0366   1.0000
   3.500   0.6668   0.02256   0.01509  -0.0583   0.0370   1.0000
   3.750   0.6953   0.02617   0.01909  -0.0581   0.0386   1.0000
   4.000   0.7210   0.03087   0.02415  -0.0578   0.0396   1.0000
   4.500   0.7790   0.03957   0.03432  -0.0540   0.0677   1.0000
   4.750   0.7986   0.04337   0.03868  -0.0532   0.0599   1.0000
   5.000   0.8188   0.04655   0.04209  -0.0528   0.0560   1.0000
   5.250   0.8370   0.05013   0.04568  -0.0527   0.0539   1.0000
   5.500   0.8444   0.05904   0.05459  -0.0532   0.0520   1.0000
   5.750   0.8633   0.06028   0.05671  -0.0521   0.0472   1.0000
   6.000   0.8753   0.06477   0.06141  -0.0524   0.0447   1.0000
   6.250   0.8856   0.06867   0.06542  -0.0525   0.0426   1.0000
   6.500   0.8962   0.07184   0.06845  -0.0516   0.0405   1.0000
   6.750   0.8022   0.07013   0.06746  -0.0435   0.0433   1.0000
   7.000   0.7965   0.07506   0.07244  -0.0439   0.0425   1.0000
   7.250   0.7790   0.08040   0.07784  -0.0450   0.0425   1.0000
   7.500   0.7562   0.08744   0.08489  -0.0505   0.0431   1.0000
<< Back to NASA SC(2)-0606 AIRFOIL (sc20606-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-0606 AIRFOIL (sc20606-il)