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)-0406 AIRFOIL (sc20406-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0406 AIRFOIL (sc20406-il)
Reynolds number: 200,000
Max Cl/Cd: 36.85 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20406-il-200000-n5.txt
Download as CSV file: xf-sc20406-il-200000-n5.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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.6885   0.08694   0.08346   0.0052   1.0000   0.0137
  -8.750  -0.6936   0.08032   0.07688  -0.0003   1.0000   0.0134
  -8.500  -0.6990   0.07089   0.06747  -0.0116   1.0000   0.0130
  -8.250  -0.7017   0.06152   0.05791  -0.0192   1.0000   0.0126
  -8.000  -0.6997   0.05260   0.04867  -0.0239   1.0000   0.0123
  -7.750  -0.6890   0.04704   0.04280  -0.0262   1.0000   0.0125
  -7.500  -0.6717   0.04468   0.04023  -0.0273   1.0000   0.0135
  -7.250  -0.6541   0.04033   0.03554  -0.0287   1.0000   0.0144
  -7.000  -0.6350   0.03458   0.02924  -0.0299   1.0000   0.0143
  -6.750  -0.6126   0.02921   0.02321  -0.0307   1.0000   0.0141
  -6.500  -0.5876   0.02523   0.01862  -0.0311   1.0000   0.0140
  -6.250  -0.5613   0.02240   0.01532  -0.0313   1.0000   0.0141
  -6.000  -0.5347   0.02024   0.01281  -0.0313   1.0000   0.0144
  -5.750  -0.5080   0.01851   0.01082  -0.0313   1.0000   0.0149
  -5.500  -0.4811   0.01709   0.00920  -0.0312   1.0000   0.0156
  -5.250  -0.4541   0.01589   0.00782  -0.0312   1.0000   0.0167
  -5.000  -0.4267   0.01493   0.00672  -0.0313   1.0000   0.0178
  -4.750  -0.3992   0.01400   0.00576  -0.0316   1.0000   0.0213
  -4.500  -0.3715   0.01345   0.00515  -0.0317   1.0000   0.0268
  -4.250  -0.3433   0.01269   0.00436  -0.0320   1.0000   0.0366
  -4.000  -0.3153   0.01221   0.00384  -0.0322   1.0000   0.0492
  -3.750  -0.2876   0.01181   0.00347  -0.0324   1.0000   0.0656
  -3.500  -0.2597   0.01133   0.00309  -0.0328   1.0000   0.0913
  -3.250  -0.2315   0.01078   0.00273  -0.0332   1.0000   0.1366
  -3.000  -0.2030   0.00999   0.00240  -0.0341   1.0000   0.2532
  -2.750  -0.1743   0.00885   0.00217  -0.0352   1.0000   0.4855
  -2.500  -0.1477   0.00848   0.00218  -0.0349   1.0000   0.5974
  -2.250  -0.1219   0.00831   0.00220  -0.0344   1.0000   0.6568
  -2.000  -0.0967   0.00821   0.00224  -0.0336   1.0000   0.7049
  -1.750  -0.0718   0.00815   0.00226  -0.0327   1.0000   0.7403
  -1.500  -0.0475   0.00812   0.00231  -0.0317   1.0000   0.7699
  -1.250  -0.0236   0.00810   0.00236  -0.0307   1.0000   0.7963
  -1.000   0.0002   0.00808   0.00241  -0.0296   1.0000   0.8148
  -0.750   0.0249   0.00807   0.00243  -0.0289   1.0000   0.8283
  -0.500   0.0498   0.00806   0.00247  -0.0283   1.0000   0.8404
  -0.250   0.0745   0.00807   0.00253  -0.0276   1.0000   0.8522
   0.000   0.0993   0.00808   0.00261  -0.0270   1.0000   0.8640
   0.250   0.1237   0.00810   0.00271  -0.0264   1.0000   0.8763
   0.500   0.1708   0.00800   0.00269  -0.0304   0.9827   0.8849
   0.750   0.2144   0.00788   0.00263  -0.0334   0.9491   0.8942
   1.000   0.2503   0.00779   0.00256  -0.0343   0.8970   0.9051
   1.250   0.2753   0.00782   0.00250  -0.0327   0.8209   0.9186
   1.500   0.2944   0.00810   0.00240  -0.0298   0.7020   0.9349
   1.750   0.3135   0.00892   0.00238  -0.0276   0.4975   0.9564
   2.000   0.3390   0.01008   0.00261  -0.0278   0.2686   1.0000
   2.250   0.3664   0.01099   0.00295  -0.0284   0.1380   1.0000
   2.500   0.3949   0.01156   0.00334  -0.0289   0.0904   1.0000
   2.750   0.4234   0.01206   0.00372  -0.0293   0.0630   1.0000
   3.000   0.4518   0.01252   0.00409  -0.0297   0.0433   1.0000
   3.250   0.4801   0.01303   0.00460  -0.0299   0.0313   1.0000
   3.500   0.5078   0.01381   0.00538  -0.0300   0.0229   1.0000
   3.750   0.5355   0.01457   0.00630  -0.0300   0.0203   1.0000
   4.000   0.5626   0.01549   0.00732  -0.0299   0.0187   1.0000
   4.250   0.5895   0.01654   0.00847  -0.0297   0.0178   1.0000
   4.500   0.6163   0.01777   0.00982  -0.0295   0.0171   1.0000
   4.750   0.6429   0.01897   0.01114  -0.0293   0.0159   1.0000
   5.000   0.6689   0.02069   0.01311  -0.0291   0.0145   1.0000
   5.250   0.6952   0.02250   0.01526  -0.0288   0.0139   1.0000
   5.500   0.7205   0.02497   0.01817  -0.0282   0.0134   1.0000
   5.750   0.7441   0.02823   0.02198  -0.0275   0.0131   1.0000
   6.000   0.7651   0.03249   0.02686  -0.0267   0.0130   1.0000
   6.250   0.7828   0.03775   0.03271  -0.0258   0.0130   1.0000
   6.500   0.7976   0.04359   0.03907  -0.0252   0.0132   1.0000
   6.750   0.8096   0.04962   0.04551  -0.0251   0.0134   1.0000
   7.000   0.8189   0.05571   0.05193  -0.0255   0.0137   1.0000
   7.250   0.8253   0.06181   0.05829  -0.0265   0.0139   1.0000
   7.500   0.8285   0.06794   0.06463  -0.0282   0.0142   1.0000
   7.750   0.8285   0.07408   0.07090  -0.0306   0.0144   1.0000
   8.000   0.8258   0.08033   0.07726  -0.0339   0.0146   1.0000
   8.250   0.7144   0.07578   0.07291  -0.0285   0.0145   1.0000
<< Back to NASA SC(2)-0406 AIRFOIL (sc20406-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-0406 AIRFOIL (sc20406-il)