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NASA SC(2)-1006 AIRFOIL (sc21006-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-1006 AIRFOIL (sc21006-il)
Reynolds number: 200,000
Max Cl/Cd: 66.54 at α=-0.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc21006-il-200000.txt
Download as CSV file: xf-sc21006-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.5059   0.09792   0.09430  -0.0227   0.9999   0.0399
 -11.000  -0.5100   0.09497   0.09140  -0.0229   0.9999   0.0414
 -10.750  -0.5182   0.09221   0.08871  -0.0229   0.9999   0.0428
 -10.500  -0.4466   0.07688   0.07368  -0.0238   0.9999   0.0479
 -10.250  -0.4511   0.07399   0.07082  -0.0226   0.9999   0.0490
 -10.000  -0.5358   0.08060   0.07729  -0.0343   0.9999   0.0446
  -9.750  -0.5285   0.07410   0.07066  -0.0435   0.9999   0.0448
  -9.500  -0.5333   0.06979   0.06651  -0.0388   0.9999   0.0467
  -9.250  -0.5275   0.06900   0.06575  -0.0346   0.9999   0.0492
  -9.000  -0.5139   0.06364   0.06033  -0.0407   0.9999   0.0515
  -8.750  -0.4722   0.05143   0.04731  -0.0629   0.9999   0.0585
  -8.500  -0.4590   0.04781   0.04389  -0.0621   0.9999   0.0604
  -8.250  -0.4369   0.04475   0.04073  -0.0640   0.9999   0.0639
  -8.000  -0.3958   0.03873   0.03417  -0.0725   0.9999   0.0738
  -7.750  -0.3298   0.02801   0.02218  -0.0822   0.9999   0.0422
  -7.500  -0.2831   0.02316   0.01661  -0.0872   0.9999   0.0386
  -7.250  -0.2441   0.02066   0.01363  -0.0898   0.9999   0.0394
  -7.000  -0.2045   0.01844   0.01101  -0.0921   0.9999   0.0392
  -6.750  -0.1681   0.01694   0.00928  -0.0939   0.9999   0.0411
  -6.500  -0.1347   0.01611   0.00830  -0.0950   0.9999   0.0448
  -6.250  -0.0941   0.01472   0.00694  -0.0981   0.9999   0.0548
  -6.000  -0.0535   0.01373   0.00597  -0.1013   0.9999   0.0743
  -5.750   0.0066   0.01170   0.00523  -0.1099   0.9999   0.3901
  -5.500   0.0397   0.01142   0.00527  -0.1113   0.9999   0.5019
  -5.250   0.0689   0.01133   0.00540  -0.1116   0.9999   0.5817
  -5.000   0.0958   0.01136   0.00556  -0.1112   0.9999   0.6363
  -4.750   0.1218   0.01145   0.00568  -0.1107   0.9999   0.6742
  -4.500   0.1473   0.01155   0.00581  -0.1101   0.9999   0.7039
  -4.250   0.1725   0.01168   0.00596  -0.1095   0.9999   0.7303
  -4.000   0.1967   0.01182   0.00614  -0.1085   0.9999   0.7567
  -3.750   0.2197   0.01198   0.00635  -0.1073   0.9999   0.7813
  -3.500   0.2428   0.01212   0.00652  -0.1061   0.9999   0.8020
  -3.250   0.2665   0.01224   0.00668  -0.1052   0.9999   0.8209
  -3.000   0.2912   0.01236   0.00683  -0.1046   0.9999   0.8391
  -2.750   0.3144   0.01244   0.00698  -0.1036   0.9999   0.8554
  -2.500   0.3377   0.01251   0.00712  -0.1026   0.9999   0.8724
  -2.250   0.3615   0.01258   0.00726  -0.1018   0.9999   0.8909
  -2.000   0.3841   0.01259   0.00736  -0.1008   0.9999   0.9104
  -1.750   0.4051   0.01257   0.00742  -0.0995   0.9999   0.9341
  -1.500   0.4152   0.01259   0.00755  -0.0962   0.9999   1.0001
  -1.250   0.4473   0.01288   0.00794  -0.0977   0.9999   1.0001
  -1.000   0.4783   0.01320   0.00834  -0.0990   0.9999   1.0001
  -0.750   0.5754   0.01204   0.00741  -0.1123   0.9672   1.0001
  -0.500   0.6494   0.00976   0.00542  -0.1188   0.9144   1.0001
  -0.250   0.6714   0.01305   0.00484  -0.1155   0.0969   1.0001
   0.000   0.6974   0.01444   0.00600  -0.1152   0.0571   1.0001
   0.250   0.7238   0.01565   0.00717  -0.1149   0.0458   1.0001
   0.500   0.7515   0.01670   0.00830  -0.1147   0.0416   1.0001
   0.750   0.7794   0.01803   0.00967  -0.1144   0.0387   1.0001
   1.000   0.8084   0.01963   0.01133  -0.1142   0.0372   1.0001
   1.250   0.8375   0.02149   0.01322  -0.1143   0.0348   1.0001
   1.500   0.8671   0.02521   0.01719  -0.1142   0.0333   1.0001
   1.750   0.8962   0.02691   0.01917  -0.1137   0.0339   1.0001
   2.000   0.9262   0.03137   0.02441  -0.1116   0.0436   1.0001
   3.250   1.0301   0.05026   0.04561  -0.1014   0.0507   1.0001
   3.500   1.0453   0.05356   0.04900  -0.1003   0.0480   1.0001
   3.750   1.0585   0.05908   0.05440  -0.0999   0.0462   1.0001
   4.000   1.0567   0.06641   0.06214  -0.0975   0.0451   1.0001
   4.250   1.0666   0.06771   0.06422  -0.0941   0.0406   1.0001
   4.500   1.0710   0.07216   0.06889  -0.0926   0.0386   1.0001
   4.750   1.0730   0.07650   0.07340  -0.0913   0.0372   1.0001
   5.000   1.0730   0.08074   0.07776  -0.0900   0.0362   1.0001
   5.250   1.0715   0.08488   0.08200  -0.0888   0.0354   1.0001
   5.500   1.0693   0.08898   0.08615  -0.0876   0.0347   1.0001
   5.750   1.0658   0.09348   0.09067  -0.0865   0.0341   1.0001
   6.000   1.0540   0.10094   0.09811  -0.0860   0.0334   1.0001
   6.250   0.8580   0.10306   0.10065  -0.0794   0.0418   1.0001
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