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NASA SC(2)-0402 AIRFOIL (sc20402-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0402 AIRFOIL (sc20402-il)
Reynolds number: 100,000
Max Cl/Cd: 16.94 at α=1.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20402-il-100000-n5.txt
Download as CSV file: xf-sc20402-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0402 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.6430   0.09175   0.08723   0.0023   1.0000   0.0289
  -7.250  -0.6472   0.08552   0.08108   0.0080   1.0000   0.0306
  -7.000  -0.6375   0.08106   0.07660   0.0061   1.0000   0.0320
  -6.750  -0.6240   0.07644   0.07193   0.0022   1.0000   0.0334
  -6.500  -0.6074   0.07167   0.06708  -0.0024   1.0000   0.0352
  -6.250  -0.5873   0.06680   0.06208  -0.0074   1.0000   0.0374
  -6.000  -0.5502   0.06310   0.05788  -0.0160   1.0000   0.0414
  -5.750  -0.5013   0.04176   0.03647  -0.0174   1.0000   0.0442
  -5.500  -0.4850   0.03713   0.03173  -0.0190   1.0000   0.0469
  -5.250  -0.4536   0.03336   0.02724  -0.0236   1.0000   0.0552
  -5.000  -0.4397   0.02805   0.02205  -0.0242   1.0000   0.0578
  -4.750  -0.4140   0.02435   0.01790  -0.0264   1.0000   0.0694
  -4.250  -0.3707   0.03200   0.02426  -0.0281   1.0000   0.0233
  -4.000  -0.3363   0.02930   0.02087  -0.0279   1.0000   0.0172
  -3.750  -0.3076   0.02635   0.01759  -0.0285   1.0000   0.0147
  -3.500  -0.2777   0.02360   0.01440  -0.0289   1.0000   0.0134
  -3.250  -0.2480   0.02133   0.01172  -0.0289   1.0000   0.0124
  -3.000  -0.2189   0.01938   0.00943  -0.0287   1.0000   0.0115
  -2.750  -0.1906   0.01771   0.00752  -0.0283   1.0000   0.0109
  -2.500  -0.1630   0.01630   0.00597  -0.0278   1.0000   0.0107
  -2.250  -0.1357   0.01519   0.00478  -0.0276   1.0000   0.0110
  -2.000  -0.1083   0.01434   0.00387  -0.0275   1.0000   0.0120
  -1.250  -0.0459   0.00905   0.00123  -0.0211   1.0000   1.0000
  -1.000  -0.0182   0.00905   0.00100  -0.0211   1.0000   1.0000
  -0.750   0.0093   0.00904   0.00086  -0.0211   1.0000   1.0000
  -0.500   0.0367   0.00905   0.00080  -0.0211   1.0000   1.0000
  -0.250   0.0640   0.00906   0.00082  -0.0211   1.0000   1.0000
   0.000   0.0912   0.00909   0.00091  -0.0211   1.0000   1.0000
   0.250   0.1183   0.00912   0.00108  -0.0210   1.0000   1.0000
   0.500   0.1455   0.00916   0.00136  -0.0208   1.0000   1.0000
   1.000   0.2212   0.01442   0.00397  -0.0238   0.0119   1.0000
   1.250   0.2480   0.01528   0.00489  -0.0234   0.0109   1.0000
   1.500   0.2748   0.01643   0.00611  -0.0230   0.0107   1.0000
   1.750   0.3024   0.01785   0.00767  -0.0225   0.0109   1.0000
   2.000   0.3307   0.01955   0.00959  -0.0220   0.0115   1.0000
   2.250   0.3594   0.02149   0.01185  -0.0214   0.0123   1.0000
   2.500   0.3881   0.02371   0.01446  -0.0209   0.0133   1.0000
   2.750   0.4162   0.02635   0.01753  -0.0206   0.0146   1.0000
   3.000   0.4427   0.02952   0.02102  -0.0205   0.0169   1.0000
   3.250   0.4754   0.03163   0.02379  -0.0192   0.0227   1.0000
   4.250   0.5873   0.04656   0.04036  -0.0195   0.0502   1.0000
   4.500   0.6111   0.05032   0.04441  -0.0205   0.0458   1.0000
   4.750   0.6285   0.05480   0.04895  -0.0211   0.0431   1.0000
   5.000   0.6419   0.06194   0.05666  -0.0222   0.0415   1.0000
   5.250   0.6777   0.06407   0.05949  -0.0266   0.0376   1.0000
   5.500   0.6982   0.06872   0.06433  -0.0297   0.0351   1.0000
   5.750   0.7154   0.07341   0.06914  -0.0326   0.0331   1.0000
   6.000   0.7291   0.07807   0.07389  -0.0349   0.0314   1.0000
   6.250   0.7373   0.08289   0.07873  -0.0353   0.0300   1.0000
   6.500   0.7355   0.09073   0.08643  -0.0338   0.0287   1.0000
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