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NASA SC(2)-0406 AIRFOIL (sc20406-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0406 AIRFOIL (sc20406-il)
Reynolds number: 500,000
Max Cl/Cd: 52.94 at α=1.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20406-il-500000.txt
Download as CSV file: xf-sc20406-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0406 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.6817   0.09078   0.08857   0.0094   1.0000   0.0179
  -8.750  -0.6824   0.08578   0.08359   0.0058   1.0000   0.0182
  -8.500  -0.6847   0.08022   0.07806   0.0009   1.0000   0.0184
  -8.250  -0.6839   0.07208   0.06990  -0.0104   1.0000   0.0185
  -8.000  -0.6794   0.06531   0.06302  -0.0170   1.0000   0.0189
  -7.750  -0.6714   0.05907   0.05664  -0.0217   1.0000   0.0195
  -6.500  -0.5850   0.02436   0.01951  -0.0315   1.0000   0.0155
  -6.250  -0.5583   0.02257   0.01748  -0.0317   1.0000   0.0152
  -6.000  -0.5319   0.01868   0.01310  -0.0323   1.0000   0.0155
  -5.750  -0.5045   0.01651   0.01066  -0.0324   1.0000   0.0156
  -5.500  -0.4769   0.01490   0.00887  -0.0324   1.0000   0.0154
  -5.250  -0.4493   0.01298   0.00674  -0.0326   1.0000   0.0158
  -5.000  -0.4210   0.01161   0.00526  -0.0329   1.0000   0.0170
  -4.750  -0.3925   0.01094   0.00455  -0.0332   1.0000   0.0189
  -4.500  -0.3639   0.01039   0.00394  -0.0335   1.0000   0.0214
  -4.250  -0.3347   0.00969   0.00322  -0.0339   1.0000   0.0287
  -4.000  -0.3061   0.00930   0.00281  -0.0341   1.0000   0.0444
  -3.750  -0.2777   0.00894   0.00248  -0.0344   1.0000   0.0566
  -3.500  -0.2494   0.00861   0.00221  -0.0347   1.0000   0.0745
  -3.250  -0.2206   0.00811   0.00194  -0.0352   1.0000   0.1248
  -3.000  -0.1903   0.00692   0.00159  -0.0368   1.0000   0.3426
  -2.750  -0.1607   0.00608   0.00146  -0.0379   1.0000   0.5437
  -2.500  -0.1329   0.00584   0.00143  -0.0379   1.0000   0.6142
  -2.250  -0.1057   0.00570   0.00144  -0.0378   1.0000   0.6630
  -2.000  -0.0787   0.00563   0.00144  -0.0375   1.0000   0.6959
  -1.750  -0.0520   0.00558   0.00145  -0.0372   1.0000   0.7225
  -1.500  -0.0256   0.00554   0.00150  -0.0368   1.0000   0.7474
  -1.250   0.0006   0.00553   0.00155  -0.0364   1.0000   0.7685
  -1.000   0.0266   0.00553   0.00163  -0.0359   1.0000   0.7909
  -0.750   0.0519   0.00554   0.00173  -0.0352   1.0000   0.8120
  -0.500   0.0776   0.00556   0.00182  -0.0347   1.0000   0.8281
  -0.250   0.1037   0.00559   0.00191  -0.0343   1.0000   0.8403
   0.000   0.1363   0.00557   0.00195  -0.0354   0.9968   0.8513
   0.250   0.1832   0.00543   0.00187  -0.0395   0.9844   0.8599
   0.500   0.2223   0.00534   0.00184  -0.0416   0.9621   0.8680
   0.750   0.2530   0.00534   0.00185  -0.0417   0.9236   0.8776
   1.000   0.2759   0.00541   0.00188  -0.0400   0.8773   0.8879
   1.250   0.2970   0.00561   0.00188  -0.0379   0.8034   0.8984
   1.500   0.3179   0.00612   0.00188  -0.0360   0.6655   0.9098
   1.750   0.3385   0.00718   0.00203  -0.0347   0.4243   0.9223
   2.000   0.3595   0.00815   0.00223  -0.0337   0.2171   0.9367
   2.250   0.3802   0.00877   0.00242  -0.0324   0.1002   0.9567
   2.500   0.4068   0.00906   0.00256  -0.0322   0.0643   1.0000
   2.750   0.4372   0.00940   0.00286  -0.0329   0.0504   1.0000
   3.000   0.4670   0.00977   0.00321  -0.0335   0.0373   1.0000
   3.250   0.4962   0.01040   0.00381  -0.0338   0.0237   1.0000
   3.500   0.5241   0.01139   0.00487  -0.0339   0.0198   1.0000
   3.750   0.5527   0.01189   0.00542  -0.0341   0.0181   1.0000
   4.000   0.5805   0.01268   0.00629  -0.0340   0.0168   1.0000
   4.250   0.6081   0.01354   0.00723  -0.0340   0.0157   1.0000
   4.500   0.6353   0.01464   0.00841  -0.0338   0.0148   1.0000
   4.750   0.6620   0.01626   0.01022  -0.0335   0.0144   1.0000
   5.000   0.6886   0.01864   0.01294  -0.0328   0.0148   1.0000
   6.500   0.7845   0.05406   0.05122  -0.0283   0.0202   1.0000
   6.750   0.7954   0.05933   0.05676  -0.0285   0.0201   1.0000
   7.000   0.8081   0.06437   0.06204  -0.0294   0.0198   1.0000
   7.250   0.8262   0.07014   0.06804  -0.0320   0.0179   1.0000
   7.500   0.8306   0.07647   0.07449  -0.0355   0.0172   1.0000
   7.750   0.8317   0.08283   0.08093  -0.0400   0.0168   1.0000
   8.000   0.8278   0.08926   0.08740  -0.0459   0.0167   1.0000
   8.250   0.8213   0.09542   0.09354  -0.0521   0.0166   1.0000
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