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NASA SC(2)-1006 AIRFOIL (sc21006-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-1006 AIRFOIL (sc21006-il)
Reynolds number: 100,000
Max Cl/Cd: 39.16 at α=0.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc21006-il-100000.txt
Download as CSV file: xf-sc21006-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.750  -0.4124   0.10007   0.09526  -0.0238   0.9999   0.0929
 -11.500  -0.4216   0.09723   0.09249  -0.0249   0.9999   0.0963
 -11.250  -0.5066   0.10306   0.09799  -0.0220   0.9999   0.0886
 -11.000  -0.5054   0.10023   0.09519  -0.0216   0.9999   0.0920
 -10.750  -0.5129   0.09782   0.09287  -0.0220   0.9999   0.0954
 -10.500  -0.5329   0.09627   0.09146  -0.0233   0.9999   0.0972
 -10.250  -0.5454   0.09215   0.08745  -0.0361   0.9999   0.0980
 -10.000  -0.5267   0.08893   0.08420  -0.0209   0.9999   0.1020
  -9.750  -0.5247   0.08598   0.08125  -0.0209   0.9999   0.1064
  -9.500  -0.5288   0.07889   0.07420  -0.0423   0.9999   0.1122
  -9.250  -0.5224   0.07735   0.07274  -0.0296   0.9999   0.1148
  -9.000  -0.5080   0.07023   0.06546  -0.0470   0.9999   0.1261
  -8.750  -0.5026   0.06822   0.06357  -0.0383   0.9999   0.1287
  -8.500  -0.4855   0.06302   0.05827  -0.0453   0.9999   0.1416
  -8.250  -0.4705   0.05960   0.05482  -0.0465   0.9999   0.1562
  -8.000  -0.4538   0.05634   0.05154  -0.0475   0.9999   0.1715
  -7.750  -0.3477   0.03600   0.02911  -0.0793   0.9999   0.0828
  -7.500  -0.2995   0.02964   0.02187  -0.0849   0.9999   0.0677
  -7.250  -0.2581   0.02587   0.01745  -0.0882   0.9999   0.0647
  -7.000  -0.2212   0.02351   0.01457  -0.0901   0.9999   0.0672
  -6.750  -0.1860   0.02160   0.01221  -0.0915   0.9999   0.0723
  -6.500  -0.1526   0.01983   0.01041  -0.0927   0.9999   0.0781
  -6.250  -0.1181   0.01847   0.00905  -0.0941   0.9999   0.0923
  -6.000  -0.0786   0.01699   0.00770  -0.0966   0.9999   0.1238
  -5.750  -0.0225   0.01437   0.00703  -0.1038   0.9999   0.5332
  -5.500  -0.0003   0.01448   0.00736  -0.1018   0.9999   0.6307
  -5.250   0.0191   0.01471   0.00763  -0.0991   0.9999   0.6870
  -5.000   0.0386   0.01488   0.00779  -0.0966   0.9999   0.7277
  -4.750   0.0567   0.01499   0.00788  -0.0938   0.9999   0.7597
  -4.500   0.0741   0.01504   0.00789  -0.0909   0.9999   0.7916
  -4.250   0.0912   0.01504   0.00787  -0.0880   0.9999   0.8256
  -4.000   0.1051   0.01488   0.00771  -0.0844   0.9999   0.8576
  -3.750   0.1207   0.01460   0.00741  -0.0815   0.9999   0.8918
  -3.500   0.1337   0.01410   0.00692  -0.0782   0.9999   0.9287
  -3.250   0.1391   0.01355   0.00642  -0.0740   0.9999   1.0001
  -3.000   0.1798   0.01373   0.00648  -0.0778   0.9999   1.0001
  -2.750   0.2178   0.01393   0.00661  -0.0808   0.9999   1.0001
  -2.500   0.2538   0.01415   0.00678  -0.0834   0.9999   1.0001
  -2.250   0.2882   0.01439   0.00701  -0.0855   0.9999   1.0001
  -2.000   0.3212   0.01464   0.00727  -0.0873   0.9999   1.0001
  -1.750   0.3531   0.01491   0.00756  -0.0887   0.9999   1.0001
  -1.500   0.3841   0.01520   0.00791  -0.0900   0.9999   1.0001
  -1.250   0.4142   0.01552   0.00833  -0.0910   0.9999   1.0001
  -1.000   0.4435   0.01585   0.00877  -0.0919   0.9999   1.0001
  -0.750   0.4722   0.01621   0.00926  -0.0926   0.9999   1.0001
  -0.500   0.5003   0.01660   0.00981  -0.0932   0.9999   1.0001
  -0.250   0.5278   0.01702   0.01042  -0.0937   0.9999   1.0001
   0.000   0.5548   0.01748   0.01116  -0.0941   0.9999   1.0001
   0.250   0.7022   0.01793   0.00809  -0.1089   0.0966   1.0001
   0.500   0.7269   0.01942   0.00957  -0.1079   0.0799   1.0001
   0.750   0.7534   0.02142   0.01139  -0.1074   0.0719   1.0001
   1.000   0.7832   0.02288   0.01297  -0.1072   0.0641   1.0001
   1.250   0.8143   0.02496   0.01509  -0.1072   0.0607   1.0001
   1.500   0.8463   0.02751   0.01783  -0.1071   0.0600   1.0001
   1.750   0.8779   0.03022   0.02111  -0.1062   0.0627   1.0001
   2.000   0.9070   0.03363   0.02514  -0.1050   0.0661   1.0001
   2.250   0.9330   0.03704   0.02902  -0.1037   0.0674   1.0001
   3.000   1.0179   0.05594   0.05094  -0.0974   0.1846   1.0001
   3.250   1.0331   0.06014   0.05506  -0.0961   0.1682   1.0001
   3.500   1.0375   0.06336   0.05866  -0.0943   0.1476   1.0001
   5.000   0.9745   0.08317   0.07940  -0.0733   0.1114   1.0001
   5.250   0.9456   0.08769   0.08416  -0.0722   0.1107   1.0001
   5.500   0.9173   0.09214   0.08871  -0.0710   0.1104   1.0001
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