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NASA SC(2)-0706 AIRFOIL (sc20706-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0706 AIRFOIL (sc20706-il)
Reynolds number: 500,000
Max Cl/Cd: 67.04 at α=0.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20706-il-500000.txt
Download as CSV file: xf-sc20706-il-500000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4914   0.08303   0.08082  -0.0184   1.0000   0.0184
  -9.750  -0.4957   0.07867   0.07648  -0.0192   1.0000   0.0187
  -9.500  -0.5016   0.07421   0.07205  -0.0200   1.0000   0.0190
  -9.250  -0.5092   0.06967   0.06753  -0.0209   1.0000   0.0193
  -9.000  -0.5196   0.06515   0.06305  -0.0216   1.0000   0.0195
  -8.750  -0.5348   0.06089   0.05882  -0.0218   1.0000   0.0194
  -8.500  -0.5539   0.05564   0.05362  -0.0246   1.0000   0.0190
  -8.250  -0.5661   0.04854   0.04643  -0.0320   1.0000   0.0188
  -8.000  -0.5679   0.04259   0.04035  -0.0358   1.0000   0.0192
  -7.750  -0.5622   0.03692   0.03451  -0.0390   1.0000   0.0200
  -7.500  -0.5420   0.03355   0.03083  -0.0407   1.0000   0.0230
  -7.250  -0.5404   0.04154   0.03820  -0.0471   1.0000   0.0234
  -6.250  -0.4015   0.01640   0.01077  -0.0631   1.0000   0.0175
  -6.000  -0.3698   0.01486   0.00901  -0.0642   1.0000   0.0182
  -5.750  -0.3388   0.01375   0.00779  -0.0650   1.0000   0.0193
  -5.500  -0.3074   0.01283   0.00677  -0.0660   1.0000   0.0205
  -5.250  -0.2773   0.01230   0.00617  -0.0666   1.0000   0.0224
  -5.000  -0.2426   0.01142   0.00518  -0.0684   1.0000   0.0243
  -4.750  -0.2046   0.01047   0.00417  -0.0710   1.0000   0.0299
  -4.500  -0.1717   0.01006   0.00369  -0.0723   1.0000   0.0360
  -4.250  -0.1383   0.00963   0.00330  -0.0738   1.0000   0.0513
  -4.000  -0.0973   0.00876   0.00292  -0.0774   1.0000   0.1729
  -3.750  -0.0458   0.00742   0.00266  -0.0840   1.0000   0.4801
  -3.500  -0.0123   0.00716   0.00271  -0.0855   1.0000   0.5844
  -3.250   0.0171   0.00711   0.00275  -0.0859   1.0000   0.6238
  -3.000   0.0456   0.00712   0.00279  -0.0860   1.0000   0.6519
  -2.750   0.0736   0.00715   0.00288  -0.0860   1.0000   0.6799
  -2.500   0.1013   0.00720   0.00296  -0.0859   1.0000   0.6998
  -2.250   0.1285   0.00726   0.00307  -0.0857   1.0000   0.7172
  -2.000   0.1558   0.00732   0.00317  -0.0855   1.0000   0.7305
  -1.750   0.1829   0.00740   0.00329  -0.0853   1.0000   0.7436
  -1.500   0.2098   0.00749   0.00342  -0.0851   1.0000   0.7553
  -1.250   0.2367   0.00756   0.00355  -0.0848   1.0000   0.7642
  -1.000   0.2640   0.00764   0.00369  -0.0848   1.0000   0.7725
  -0.750   0.2910   0.00772   0.00383  -0.0846   1.0000   0.7797
  -0.500   0.3181   0.00781   0.00399  -0.0845   1.0000   0.7869
  -0.250   0.3902   0.00707   0.00336  -0.0939   0.9811   0.7922
   0.000   0.4355   0.00666   0.00304  -0.0973   0.9645   0.7986
   0.250   0.4887   0.00729   0.00246  -0.1016   0.6233   0.8022
   0.500   0.5047   0.00914   0.00289  -0.0997   0.2786   0.8083
   0.750   0.5269   0.01046   0.00330  -0.0990   0.0670   0.8142
   1.000   0.5527   0.01098   0.00377  -0.0984   0.0419   0.8198
   1.250   0.5796   0.01146   0.00426  -0.0981   0.0316   0.8263
   1.500   0.6053   0.01189   0.00477  -0.0974   0.0274   0.8316
   1.750   0.6317   0.01241   0.00534  -0.0970   0.0239   0.8382
   2.000   0.6549   0.01368   0.00670  -0.0957   0.0211   0.8436
   2.250   0.6792   0.01508   0.00825  -0.0946   0.0202   0.8497
   2.500   0.7060   0.01556   0.00882  -0.0942   0.0191   0.8554
   2.750   0.7314   0.01651   0.00990  -0.0934   0.0181   0.8610
   3.000   0.7579   0.01790   0.01145  -0.0927   0.0174   0.8673
   3.250   0.7827   0.01961   0.01342  -0.0915   0.0169   0.8722
   3.500   0.8073   0.02290   0.01718  -0.0898   0.0175   0.8779
   4.750   0.8861   0.04652   0.04289  -0.0787   0.0205   0.9074
   5.000   0.8980   0.05035   0.04706  -0.0764   0.0204   0.9138
   5.250   0.9133   0.05360   0.05066  -0.0743   0.0201   0.9214
   5.500   0.9359   0.05633   0.05370  -0.0723   0.0172   0.9294
   5.750   0.9442   0.06071   0.05831  -0.0708   0.0163   0.9385
   6.000   0.9499   0.06489   0.06269  -0.0694   0.0156   0.9504
   6.250   0.9527   0.06901   0.06698  -0.0681   0.0152   0.9999
   6.500   0.9586   0.07385   0.07195  -0.0687   0.0148   1.0000
   6.750   0.9615   0.07880   0.07702  -0.0696   0.0146   1.0000
   7.000   0.9615   0.08383   0.08214  -0.0708   0.0144   1.0000
   7.250   0.9581   0.08918   0.08756  -0.0726   0.0142   1.0000
   7.500   0.9482   0.09456   0.09299  -0.0746   0.0142   1.0000
   7.750   0.9373   0.10160   0.10005  -0.0810   0.0142   1.0000
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