Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NASA SC(2)-0706 AIRFOIL (sc20706-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0706 AIRFOIL (sc20706-il)
Reynolds number: 500,000
Max Cl/Cd: 55.8 at α=-0.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20706-il-500000-n5.txt
Download as CSV file: xf-sc20706-il-500000-n5.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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.6307   0.08261   0.08031  -0.0175   1.0000   0.0070
  -9.750  -0.6426   0.07737   0.07515  -0.0193   1.0000   0.0069
  -9.500  -0.6583   0.07190   0.06974  -0.0215   1.0000   0.0069
  -9.250  -0.6719   0.06013   0.05791  -0.0347   1.0000   0.0068
  -9.000  -0.6762   0.05089   0.04845  -0.0412   1.0000   0.0068
  -8.750  -0.6702   0.04177   0.03895  -0.0465   1.0000   0.0069
  -8.500  -0.6524   0.03394   0.03063  -0.0509   1.0000   0.0072
  -8.250  -0.6262   0.02790   0.02403  -0.0547   1.0000   0.0075
  -8.000  -0.5991   0.02536   0.02118  -0.0565   1.0000   0.0079
  -7.750  -0.5707   0.02301   0.01847  -0.0581   1.0000   0.0085
  -7.500  -0.5382   0.01962   0.01457  -0.0605   1.0000   0.0087
  -7.250  -0.5067   0.01732   0.01191  -0.0621   1.0000   0.0090
  -7.000  -0.4760   0.01575   0.01007  -0.0632   1.0000   0.0094
  -6.750  -0.4458   0.01460   0.00873  -0.0642   1.0000   0.0099
  -6.500  -0.4159   0.01377   0.00776  -0.0650   1.0000   0.0104
  -6.250  -0.3865   0.01312   0.00698  -0.0657   1.0000   0.0107
  -6.000  -0.3524   0.01199   0.00572  -0.0676   1.0000   0.0116
  -5.750  -0.3218   0.01148   0.00517  -0.0685   1.0000   0.0127
  -5.500  -0.2924   0.01117   0.00483  -0.0692   1.0000   0.0142
  -5.250  -0.2613   0.01077   0.00438  -0.0701   1.0000   0.0157
  -5.000  -0.2308   0.01048   0.00400  -0.0709   1.0000   0.0172
  -4.750  -0.1983   0.01006   0.00357  -0.0722   1.0000   0.0222
  -4.500  -0.1686   0.00988   0.00335  -0.0728   1.0000   0.0263
  -4.250  -0.1382   0.00967   0.00316  -0.0735   1.0000   0.0347
  -4.000  -0.1071   0.00942   0.00299  -0.0745   1.0000   0.0566
  -3.750  -0.0752   0.00914   0.00286  -0.0757   1.0000   0.0944
  -3.500  -0.0417   0.00879   0.00274  -0.0774   1.0000   0.1629
  -3.250  -0.0050   0.00829   0.00263  -0.0799   1.0000   0.2766
  -3.000   0.0378   0.00752   0.00253  -0.0841   1.0000   0.4768
  -2.750   0.0708   0.00731   0.00260  -0.0855   1.0000   0.5645
  -2.500   0.1006   0.00726   0.00267  -0.0860   1.0000   0.6053
  -2.250   0.1296   0.00725   0.00276  -0.0862   1.0000   0.6368
  -2.000   0.1578   0.00728   0.00285  -0.0863   1.0000   0.6580
  -1.750   0.1859   0.00732   0.00294  -0.0864   1.0000   0.6728
  -1.500   0.2346   0.00708   0.00276  -0.0910   0.9935   0.6871
  -1.250   0.2774   0.00683   0.00258  -0.0942   0.9857   0.6999
  -1.000   0.3152   0.00663   0.00245  -0.0963   0.9774   0.7096
  -0.500   0.4068   0.00729   0.00175  -0.1030   0.5863   0.7244
  -0.250   0.4281   0.00842   0.00203  -0.1020   0.3657   0.7308
   0.000   0.4523   0.00925   0.00226  -0.1015   0.2037   0.7374
   0.250   0.4783   0.00980   0.00248  -0.1013   0.1124   0.7439
   0.500   0.5050   0.01018   0.00270  -0.1011   0.0662   0.7500
   1.000   0.5591   0.01078   0.00318  -0.1008   0.0279   0.7625
   1.250   0.5865   0.01104   0.00346  -0.1006   0.0230   0.7688
   1.500   0.6136   0.01131   0.00376  -0.1004   0.0189   0.7740
   1.750   0.6401   0.01177   0.00428  -0.1000   0.0153   0.7798
   2.000   0.6668   0.01215   0.00475  -0.0997   0.0140   0.7857
   2.250   0.6931   0.01263   0.00532  -0.0992   0.0130   0.7917
   2.750   0.7453   0.01365   0.00648  -0.0983   0.0114   0.8033
   3.000   0.7714   0.01415   0.00702  -0.0979   0.0105   0.8092
   3.250   0.7961   0.01511   0.00810  -0.0971   0.0099   0.8141
   3.500   0.8202   0.01646   0.00965  -0.0962   0.0095   0.8193
   4.000   0.8709   0.01838   0.01193  -0.0949   0.0089   0.8298
   4.250   0.8956   0.01982   0.01362  -0.0940   0.0085   0.8353
   4.500   0.9193   0.02180   0.01593  -0.0929   0.0082   0.8402
   4.750   0.9409   0.02463   0.01921  -0.0913   0.0079   0.8451
   5.000   0.9598   0.02854   0.02368  -0.0892   0.0077   0.8505
   5.250   0.9780   0.03182   0.02740  -0.0873   0.0072   0.8554
   5.500   0.9941   0.03557   0.03155  -0.0853   0.0068   0.8608
   5.750   1.0046   0.04157   0.03804  -0.0825   0.0066   0.8660
   6.000   1.0145   0.04667   0.04352  -0.0802   0.0064   0.8708
   6.250   1.0219   0.05214   0.04931  -0.0782   0.0062   0.8763
   6.500   1.0269   0.05759   0.05504  -0.0765   0.0060   0.8817
   6.750   1.0255   0.06433   0.06206  -0.0752   0.0059   0.8875
   7.000   1.0186   0.07183   0.06980  -0.0746   0.0058   0.8938
   7.250   1.0101   0.07885   0.07700  -0.0750   0.0058   0.9008
   7.500   0.9996   0.08560   0.08387  -0.0765   0.0058   0.9089
<< Back to NASA SC(2)-0706 AIRFOIL (sc20706-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-0706 AIRFOIL (sc20706-il)