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NASA SC(2)-0606 AIRFOIL (sc20606-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0606 AIRFOIL (sc20606-il)
Reynolds number: 50,000
Max Cl/Cd: 23.34 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20606-il-50000.txt
Download as CSV file: xf-sc20606-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0606 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5835   0.09516   0.08818   0.0111   1.0000   0.3326
  -8.000  -0.5736   0.09135   0.08440   0.0124   1.0000   0.3517
  -7.750  -0.5777   0.08927   0.08240   0.0139   1.0000   0.3755
  -7.500  -0.5730   0.08651   0.07969   0.0163   1.0000   0.4039
  -7.250  -0.5430   0.08194   0.07506   0.0190   1.0000   0.4406
  -7.000  -0.5331   0.07876   0.07191   0.0210   1.0000   0.4680
  -6.500  -0.5194   0.04717   0.03879  -0.0469   1.0000   0.1359
  -6.250  -0.4854   0.04166   0.03229  -0.0510   1.0000   0.1237
  -6.000  -0.4556   0.03761   0.02773  -0.0527   1.0000   0.1230
  -5.750  -0.4247   0.03387   0.02348  -0.0538   1.0000   0.1210
  -5.500  -0.3933   0.03066   0.01973  -0.0544   1.0000   0.1205
  -5.250  -0.3621   0.02813   0.01660  -0.0546   1.0000   0.1251
  -5.000  -0.3349   0.02603   0.01431  -0.0541   1.0000   0.1385
  -4.750  -0.3101   0.02394   0.01224  -0.0526   1.0000   0.1510
  -4.500  -0.2863   0.02220   0.01055  -0.0511   1.0000   0.1807
  -4.250  -0.2607   0.02007   0.00882  -0.0503   1.0000   0.2351
  -4.000  -0.2579   0.01753   0.00923  -0.0416   1.0000   0.7054
  -3.750  -0.2689   0.01815   0.00986  -0.0287   1.0000   0.8026
  -3.500  -0.2760   0.01802   0.00965  -0.0180   1.0000   0.8668
  -3.250  -0.1719   0.01696   0.00759  -0.0264   1.0000   1.0000
  -3.000  -0.1616   0.01649   0.00696  -0.0247   1.0000   1.0000
  -2.750  -0.1512   0.01607   0.00641  -0.0229   1.0000   1.0000
  -2.500  -0.1404   0.01570   0.00592  -0.0211   1.0000   1.0000
  -2.250  -0.1252   0.01540   0.00548  -0.0200   1.0000   1.0000
  -2.000  -0.1039   0.01519   0.00513  -0.0200   1.0000   1.0000
  -1.750  -0.0785   0.01506   0.00482  -0.0208   1.0000   1.0000
  -1.500  -0.0504   0.01498   0.00461  -0.0219   1.0000   1.0000
  -1.250  -0.0210   0.01495   0.00448  -0.0232   1.0000   1.0000
  -1.000   0.0092   0.01496   0.00441  -0.0246   1.0000   1.0000
  -0.750   0.0397   0.01500   0.00439  -0.0259   1.0000   1.0000
  -0.500   0.0703   0.01507   0.00443  -0.0272   1.0000   1.0000
  -0.250   0.1008   0.01516   0.00453  -0.0285   1.0000   1.0000
   0.000   0.1311   0.01528   0.00469  -0.0297   1.0000   1.0000
   0.250   0.1612   0.01542   0.00491  -0.0308   1.0000   1.0000
   0.500   0.1911   0.01558   0.00517  -0.0318   1.0000   1.0000
   0.750   0.2206   0.01577   0.00549  -0.0327   1.0000   1.0000
   1.000   0.2498   0.01598   0.00587  -0.0336   1.0000   1.0000
   1.250   0.2787   0.01622   0.00635  -0.0343   1.0000   1.0000
   1.500   0.3072   0.01649   0.00686  -0.0351   1.0000   1.0000
   1.750   0.3354   0.01679   0.00745  -0.0357   1.0000   1.0000
   2.000   0.3633   0.01713   0.00814  -0.0363   1.0000   1.0000
   2.250   0.3909   0.01750   0.00900  -0.0369   1.0000   1.0000
   2.500   0.4185   0.01793   0.01000  -0.0373   1.0000   1.0000
   2.750   0.5413   0.02386   0.01241  -0.0462   0.1559   1.0000
   3.000   0.5757   0.02630   0.01472  -0.0467   0.1340   1.0000
   3.250   0.6094   0.02866   0.01754  -0.0466   0.1255   1.0000
   3.500   0.6395   0.03125   0.02041  -0.0465   0.1161   1.0000
   3.750   0.6701   0.03424   0.02391  -0.0460   0.1150   1.0000
   4.000   0.6988   0.03775   0.02791  -0.0456   0.1175   1.0000
   4.250   0.7272   0.04151   0.03255  -0.0448   0.1258   1.0000
   4.500   0.7524   0.04604   0.03744  -0.0445   0.1325   1.0000
   4.750   0.7783   0.05087   0.04298  -0.0442   0.1484   1.0000
   6.000   0.8155   0.09017   0.08490  -0.0920   0.4246   1.0000
   6.250   0.8241   0.09349   0.08819  -0.0878   0.3815   1.0000
   6.500   0.8275   0.09706   0.09172  -0.0856   0.3497   1.0000
   6.750   0.6612   0.09245   0.08725  -0.0703   0.3779   1.0000
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