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NASA SC(2)-0706 AIRFOIL (sc20706-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0706 AIRFOIL (sc20706-il)
Reynolds number: 1,000,000
Max Cl/Cd: 96.57 at α=-0.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20706-il-1000000.txt
Download as CSV file: xf-sc20706-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0706 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
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.5030   0.08858   0.08700  -0.0175   1.0000   0.0103
 -10.500  -0.5090   0.08370   0.08213  -0.0184   1.0000   0.0104
 -10.250  -0.5150   0.07897   0.07741  -0.0193   1.0000   0.0105
 -10.000  -0.5213   0.07438   0.07285  -0.0201   1.0000   0.0106
  -9.750  -0.5295   0.06954   0.06802  -0.0211   1.0000   0.0106
  -9.500  -0.5405   0.06454   0.06305  -0.0220   1.0000   0.0106
  -9.250  -0.5570   0.05919   0.05772  -0.0231   1.0000   0.0106
  -9.000  -0.5861   0.05251   0.05109  -0.0258   1.0000   0.0104
  -8.750  -0.6005   0.04536   0.04384  -0.0328   1.0000   0.0103
  -8.500  -0.6085   0.03857   0.03690  -0.0365   1.0000   0.0104
  -8.250  -0.6076   0.03224   0.03036  -0.0397   1.0000   0.0106
  -8.000  -0.5968   0.02703   0.02493  -0.0424   1.0000   0.0108
  -7.750  -0.5903   0.03040   0.02759  -0.0511   1.0000   0.0092
  -7.500  -0.5512   0.02173   0.01808  -0.0572   1.0000   0.0094
  -7.250  -0.5205   0.01955   0.01557  -0.0587   1.0000   0.0097
  -6.750  -0.4531   0.01457   0.00988  -0.0630   1.0000   0.0101
  -6.500  -0.4158   0.01215   0.00720  -0.0658   1.0000   0.0109
  -6.250  -0.3835   0.01130   0.00627  -0.0671   1.0000   0.0117
  -6.000  -0.3527   0.01083   0.00574  -0.0680   1.0000   0.0126
  -5.750  -0.3208   0.01032   0.00518  -0.0692   1.0000   0.0136
  -5.500  -0.2891   0.00990   0.00469  -0.0703   1.0000   0.0144
  -5.250  -0.2582   0.00958   0.00434  -0.0711   1.0000   0.0151
  -5.000  -0.2212   0.00888   0.00352  -0.0734   1.0000   0.0186
  -4.750  -0.1910   0.00868   0.00332  -0.0741   1.0000   0.0223
  -4.500  -0.1591   0.00840   0.00301  -0.0752   1.0000   0.0286
  -4.250  -0.1293   0.00826   0.00287  -0.0757   1.0000   0.0343
  -4.000  -0.0971   0.00799   0.00270  -0.0769   1.0000   0.0608
  -3.750  -0.0601   0.00749   0.00255  -0.0795   1.0000   0.1565
  -3.500  -0.0167   0.00673   0.00240  -0.0837   1.0000   0.3327
  -3.250   0.0249   0.00614   0.00232  -0.0874   1.0000   0.4949
  -3.000   0.0579   0.00597   0.00235  -0.0887   1.0000   0.5672
  -2.750   0.0877   0.00594   0.00239  -0.0892   1.0000   0.5990
  -2.500   0.1170   0.00592   0.00245  -0.0895   1.0000   0.6248
  -2.250   0.1460   0.00593   0.00254  -0.0898   1.0000   0.6530
  -2.000   0.1742   0.00597   0.00261  -0.0899   1.0000   0.6681
  -1.750   0.2023   0.00601   0.00271  -0.0900   1.0000   0.6850
  -1.500   0.2479   0.00579   0.00254  -0.0939   0.9963   0.6974
  -1.250   0.2996   0.00535   0.00217  -0.0991   0.9882   0.7112
  -1.000   0.3480   0.00492   0.00180  -0.1035   0.9784   0.7207
  -0.750   0.3901   0.00465   0.00157  -0.1064   0.9671   0.7296
  -0.500   0.4307   0.00446   0.00139  -0.1090   0.9297   0.7364
  -0.250   0.4462   0.00618   0.00158  -0.1059   0.5507   0.7429
   0.000   0.4683   0.00731   0.00184  -0.1051   0.3159   0.7492
   0.250   0.4926   0.00822   0.00208  -0.1047   0.1386   0.7562
   0.500   0.5188   0.00873   0.00230  -0.1044   0.0586   0.7620
   0.750   0.5460   0.00905   0.00252  -0.1043   0.0345   0.7684
   1.000   0.5735   0.00926   0.00277  -0.1041   0.0280   0.7739
   1.250   0.6004   0.00970   0.00324  -0.1038   0.0204   0.7797
   1.500   0.6280   0.00989   0.00346  -0.1036   0.0184   0.7856
   1.750   0.6551   0.01019   0.00379  -0.1034   0.0160   0.7911
   2.000   0.6799   0.01124   0.00499  -0.1025   0.0133   0.7969
   2.250   0.7068   0.01155   0.00535  -0.1022   0.0129   0.8023
   2.500   0.7330   0.01209   0.00597  -0.1017   0.0125   0.8084
   2.750   0.7591   0.01272   0.00667  -0.1012   0.0120   0.8142
   3.000   0.7849   0.01339   0.00744  -0.1006   0.0114   0.8195
   3.250   0.8111   0.01409   0.00824  -0.1001   0.0107   0.8252
   3.500   0.8368   0.01484   0.00909  -0.0995   0.0100   0.8301
   3.750   0.8621   0.01597   0.01035  -0.0988   0.0096   0.8354
   4.000   0.8870   0.01729   0.01183  -0.0980   0.0092   0.8406
   4.250   0.9091   0.01996   0.01483  -0.0966   0.0088   0.8451
   4.500   0.9312   0.02249   0.01771  -0.0952   0.0088   0.8503
   4.750   0.9519   0.02528   0.02089  -0.0936   0.0088   0.8552
   5.000   0.9700   0.02872   0.02475  -0.0914   0.0089   0.8601
  11.000   0.9387   0.17193   0.17071  -0.1224   0.0075   1.0000
  11.250   0.9432   0.17649   0.17526  -0.1245   0.0069   1.0000
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