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NASA SC(2)-1006 AIRFOIL (sc21006-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-1006 AIRFOIL (sc21006-il)
Reynolds number: 500,000
Max Cl/Cd: 101.92 at α=-1.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc21006-il-500000.txt
Download as CSV file: xf-sc21006-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.5375   0.08798   0.08580  -0.0190   0.9999   0.0166
 -10.500  -0.5491   0.08573   0.08358  -0.0169   0.9999   0.0168
 -10.250  -0.5595   0.08316   0.08106  -0.0156   0.9999   0.0170
 -10.000  -0.5624   0.07955   0.07748  -0.0168   0.9999   0.0173
  -9.750  -0.5621   0.07525   0.07320  -0.0199   0.9999   0.0177
  -9.500  -0.5563   0.06965   0.06762  -0.0262   0.9999   0.0181
  -9.250  -0.5405   0.06129   0.05920  -0.0381   0.9999   0.0187
  -9.000  -0.5136   0.05169   0.04937  -0.0502   0.9999   0.0205
  -8.750  -0.4786   0.04633   0.04363  -0.0568   0.9999   0.0220
  -7.750  -0.2924   0.01759   0.01235  -0.0897   0.9999   0.0179
  -7.500  -0.2535   0.01576   0.01023  -0.0923   0.9999   0.0186
  -7.250  -0.2175   0.01455   0.00885  -0.0943   0.9999   0.0196
  -7.000  -0.1856   0.01395   0.00815  -0.0954   0.9999   0.0211
  -6.750  -0.1536   0.01343   0.00752  -0.0965   0.9999   0.0220
  -6.500  -0.1046   0.01180   0.00573  -0.1015   0.9999   0.0260
  -6.250  -0.0712   0.01143   0.00531  -0.1030   0.9999   0.0306
  -6.000  -0.0362   0.01102   0.00487  -0.1048   0.9999   0.0413
  -5.750   0.0041   0.01038   0.00450  -0.1080   0.9999   0.1174
  -5.500   0.0474   0.00966   0.00436  -0.1122   0.9999   0.2819
  -5.250   0.0833   0.00936   0.00437  -0.1144   0.9999   0.3871
  -5.000   0.1167   0.00919   0.00442  -0.1159   0.9999   0.4613
  -4.750   0.1482   0.00911   0.00450  -0.1170   0.9999   0.5218
  -4.500   0.1791   0.00907   0.00462  -0.1178   0.9999   0.5789
  -4.250   0.2082   0.00911   0.00476  -0.1182   0.9999   0.6168
  -4.000   0.2365   0.00918   0.00491  -0.1184   0.9999   0.6454
  -3.750   0.2642   0.00929   0.00507  -0.1185   0.9999   0.6720
  -3.500   0.2916   0.00941   0.00526  -0.1185   0.9999   0.6954
  -3.250   0.3192   0.00954   0.00546  -0.1186   0.9997   0.7190
  -3.000   0.3545   0.00956   0.00556  -0.1202   0.9974   0.7417
  -2.750   0.3992   0.00948   0.00555  -0.1238   0.9946   0.7580
  -2.500   0.4440   0.00921   0.00535  -0.1273   0.9877   0.7717
  -2.250   0.5023   0.00856   0.00477  -0.1334   0.9774   0.7838
  -2.000   0.5752   0.00713   0.00346  -0.1419   0.9605   0.7935
  -1.750   0.6138   0.00654   0.00293  -0.1433   0.9430   0.8032
  -1.500   0.6418   0.00633   0.00277  -0.1426   0.9191   0.8132
  -1.250   0.6645   0.00652   0.00243  -0.1401   0.7541   0.8217
  -1.000   0.6696   0.00910   0.00313  -0.1356   0.3517   0.8299
  -0.750   0.6866   0.01097   0.00375  -0.1340   0.0680   0.8394
  -0.500   0.7103   0.01169   0.00435  -0.1329   0.0331   0.8481
  -0.250   0.7361   0.01204   0.00477  -0.1322   0.0282   0.8566
   0.000   0.7604   0.01280   0.00561  -0.1312   0.0237   0.8654
   0.250   0.7831   0.01374   0.00667  -0.1298   0.0219   0.8731
   0.500   0.8083   0.01418   0.00717  -0.1291   0.0201   0.8821
   0.750   0.8322   0.01497   0.00805  -0.1279   0.0190   0.8917
   1.000   0.8556   0.01585   0.00903  -0.1267   0.0179   0.9016
   1.250   0.8789   0.01686   0.01015  -0.1254   0.0170   0.9128
   1.500   0.9021   0.01821   0.01167  -0.1239   0.0166   0.9253
   1.750   0.9256   0.02091   0.01472  -0.1218   0.0177   0.9410
   2.000   0.9480   0.02254   0.01660  -0.1201   0.0172   1.0001
   4.750   1.0361   0.05622   0.05389  -0.0854   0.0154   1.0001
   5.000   1.0280   0.06136   0.05917  -0.0826   0.0151   1.0001
   5.250   1.0104   0.06730   0.06527  -0.0793   0.0148   1.0001
   5.500   0.9875   0.07234   0.07043  -0.0754   0.0147   1.0001
   5.750   0.9691   0.07568   0.07388  -0.0720   0.0148   1.0001
   6.000   0.9463   0.08006   0.07837  -0.0699   0.0149   1.0001
   6.250   0.9237   0.08547   0.08388  -0.0700   0.0150   1.0001
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