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NASA SC(2)-0606 AIRFOIL (sc20606-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0606 AIRFOIL (sc20606-il)
Reynolds number: 1,000,000
Max Cl/Cd: 82.09 at α=-0.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20606-il-1000000.txt
Download as CSV file: xf-sc20606-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.5077   0.10055   0.09893  -0.0118   1.0000   0.0099
 -11.000  -0.5077   0.09669   0.09507  -0.0127   1.0000   0.0101
 -10.750  -0.6146   0.10655   0.10484  -0.0020   1.0000   0.0093
 -10.500  -0.6132   0.10197   0.10027  -0.0041   1.0000   0.0093
 -10.250  -0.6122   0.09728   0.09559  -0.0063   1.0000   0.0094
 -10.000  -0.6275   0.08892   0.08728  -0.0105   1.0000   0.0096
  -9.750  -0.6328   0.08337   0.08174  -0.0137   1.0000   0.0098
  -9.500  -0.6373   0.07817   0.07657  -0.0169   1.0000   0.0099
  -9.250  -0.6465   0.07251   0.07095  -0.0206   1.0000   0.0099
  -9.000  -0.6510   0.06240   0.06079  -0.0334   1.0000   0.0098
  -8.750  -0.6532   0.05559   0.05385  -0.0381   1.0000   0.0099
  -8.500  -0.6491   0.05018   0.04829  -0.0410   1.0000   0.0101
  -8.250  -0.6388   0.04607   0.04403  -0.0429   1.0000   0.0103
  -8.000  -0.6249   0.04172   0.03949  -0.0447   1.0000   0.0106
  -7.750  -0.6047   0.03030   0.02744  -0.0483   1.0000   0.0091
  -7.500  -0.5755   0.02292   0.01928  -0.0521   1.0000   0.0095
  -7.250  -0.5447   0.01862   0.01446  -0.0546   1.0000   0.0096
  -7.000  -0.5144   0.01577   0.01122  -0.0561   1.0000   0.0097
  -6.750  -0.4840   0.01348   0.00865  -0.0573   1.0000   0.0100
  -6.500  -0.4552   0.01234   0.00741  -0.0580   1.0000   0.0107
  -6.250  -0.4262   0.01142   0.00639  -0.0586   1.0000   0.0111
  -6.000  -0.3975   0.01075   0.00565  -0.0590   1.0000   0.0119
  -5.750  -0.3688   0.01019   0.00502  -0.0594   1.0000   0.0127
  -5.500  -0.3398   0.00969   0.00446  -0.0599   1.0000   0.0134
  -5.250  -0.3110   0.00928   0.00399  -0.0603   1.0000   0.0140
  -5.000  -0.2823   0.00893   0.00360  -0.0607   1.0000   0.0145
  -4.750  -0.2502   0.00828   0.00286  -0.0618   1.0000   0.0182
  -4.500  -0.2220   0.00809   0.00264  -0.0620   1.0000   0.0214
  -4.250  -0.1921   0.00778   0.00235  -0.0626   1.0000   0.0331
  -4.000  -0.1629   0.00757   0.00219  -0.0631   1.0000   0.0484
  -3.750  -0.1334   0.00737   0.00208  -0.0637   1.0000   0.0709
  -3.500  -0.1002   0.00693   0.00193  -0.0653   1.0000   0.1499
  -3.250  -0.0604   0.00612   0.00176  -0.0687   1.0000   0.3339
  -3.000  -0.0222   0.00548   0.00167  -0.0716   1.0000   0.5034
  -2.750   0.0090   0.00532   0.00168  -0.0725   1.0000   0.5697
  -2.500   0.0386   0.00525   0.00171  -0.0730   1.0000   0.6051
  -2.250   0.0674   0.00523   0.00175  -0.0733   1.0000   0.6288
  -2.000   0.0961   0.00522   0.00181  -0.0735   1.0000   0.6526
  -1.750   0.1246   0.00522   0.00189  -0.0736   1.0000   0.6771
  -1.500   0.1527   0.00525   0.00198  -0.0737   1.0000   0.6954
  -1.250   0.1895   0.00515   0.00192  -0.0757   0.9975   0.7097
  -1.000   0.2388   0.00481   0.00165  -0.0804   0.9892   0.7253
  -0.750   0.2849   0.00454   0.00145  -0.0843   0.9795   0.7368
  -0.500   0.3209   0.00442   0.00138  -0.0859   0.9655   0.7457
  -0.250   0.3579   0.00436   0.00131  -0.0877   0.9378   0.7544
   0.000   0.3812   0.00478   0.00131  -0.0861   0.8053   0.7617
   0.250   0.4007   0.00613   0.00154  -0.0844   0.5250   0.7690
   0.500   0.4239   0.00732   0.00179  -0.0840   0.2688   0.7753
   1.000   0.4767   0.00844   0.00219  -0.0837   0.0700   0.7879
   1.250   0.5042   0.00872   0.00238  -0.0837   0.0456   0.7943
   1.500   0.5316   0.00897   0.00259  -0.0835   0.0306   0.8002
   1.750   0.5590   0.00933   0.00295  -0.0833   0.0196   0.8066
   2.000   0.5864   0.00961   0.00328  -0.0831   0.0174   0.8126
   2.250   0.6137   0.00990   0.00364  -0.0829   0.0154   0.8188
   2.500   0.6406   0.01037   0.00416  -0.0826   0.0136   0.8252
   2.750   0.6655   0.01142   0.00537  -0.0818   0.0123   0.8311
   3.000   0.6923   0.01194   0.00596  -0.0815   0.0120   0.8377
   3.250   0.7183   0.01255   0.00666  -0.0810   0.0117   0.8433
   3.500   0.7445   0.01325   0.00747  -0.0805   0.0113   0.8497
   3.750   0.7704   0.01406   0.00839  -0.0799   0.0107   0.8554
   4.000   0.7962   0.01499   0.00944  -0.0793   0.0101   0.8617
   4.250   0.8214   0.01638   0.01102  -0.0785   0.0097   0.8677
   4.500   0.8462   0.01779   0.01264  -0.0777   0.0094   0.8736
   4.750   0.8715   0.01852   0.01349  -0.0773   0.0089   0.8798
   5.000   0.8946   0.02036   0.01560  -0.0763   0.0087   0.8857
   5.750   0.9220   0.04687   0.04441  -0.0651   0.0096   0.9042
   6.000   0.9342   0.05102   0.04878  -0.0638   0.0092   0.9115
   9.250   0.7425   0.11824   0.11706  -0.0708   0.0089   1.0000
   9.500   0.7392   0.12273   0.12155  -0.0722   0.0088   1.0000
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