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NASA/LANGLEY MS(1)-0313 AIRFOIL (ms313-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NASA/LANGLEY MS(1)-0313 AIRFOIL (ms313-il)
Reynolds number: 50,000
Max Cl/Cd: 31.58 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ms313-il-50000.txt
Download as CSV file: xf-ms313-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/LANGLEY MS(1)-0313 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4055   0.10478   0.09770  -0.0144   1.0000   0.3823
  -8.750  -0.3771   0.09970   0.09258  -0.0143   1.0000   0.3893
  -8.500  -0.5275   0.08083   0.07417  -0.0376   1.0000   0.1948
  -8.250  -0.5787   0.07408   0.06746  -0.0398   1.0000   0.1765
  -8.000  -0.6152   0.06818   0.06139  -0.0402   1.0000   0.1625
  -7.750  -0.6402   0.06218   0.05492  -0.0407   1.0000   0.1502
  -7.500  -0.6393   0.05778   0.05029  -0.0399   1.0000   0.1448
  -7.250  -0.6406   0.05254   0.04411  -0.0404   1.0000   0.1364
  -7.000  -0.6277   0.04894   0.04022  -0.0398   1.0000   0.1352
  -6.750  -0.6127   0.04567   0.03646  -0.0395   1.0000   0.1357
  -6.500  -0.5940   0.04281   0.03303  -0.0393   1.0000   0.1371
  -6.250  -0.5733   0.03984   0.02980  -0.0390   1.0000   0.1388
  -6.000  -0.5518   0.03752   0.02738  -0.0385   1.0000   0.1421
  -5.750  -0.5289   0.03566   0.02526  -0.0380   1.0000   0.1484
  -5.500  -0.5041   0.03376   0.02295  -0.0376   1.0000   0.1542
  -5.250  -0.4809   0.03218   0.02142  -0.0367   1.0000   0.1631
  -5.000  -0.4568   0.03075   0.01995  -0.0358   1.0000   0.1745
  -4.750  -0.4333   0.02952   0.01876  -0.0346   1.0000   0.1897
  -4.500  -0.4106   0.02838   0.01775  -0.0332   1.0000   0.2120
  -4.250  -0.3873   0.02700   0.01667  -0.0323   1.0000   0.2512
  -4.000  -0.3684   0.02389   0.01616  -0.0304   1.0000   0.4985
  -3.750  -0.3796   0.02630   0.01900  -0.0166   1.0000   0.6971
  -3.500  -0.3829   0.02783   0.02041  -0.0063   1.0000   0.7519
  -3.250  -0.3831   0.02865   0.02111   0.0023   1.0000   0.7948
  -3.000  -0.3825   0.02901   0.02135   0.0105   1.0000   0.8323
  -2.750  -0.3780   0.02900   0.02119   0.0172   1.0000   0.8717
  -2.500  -0.1211   0.03052   0.02157  -0.0162   1.0000   0.9883
  -2.250  -0.0747   0.03013   0.02095  -0.0220   1.0000   1.0000
  -2.000  -0.0745   0.02985   0.02061  -0.0194   1.0000   1.0000
  -1.750  -0.0740   0.02959   0.02029  -0.0167   1.0000   1.0000
  -1.500  -0.0735   0.02934   0.01998  -0.0140   1.0000   1.0000
  -1.250  -0.0728   0.02910   0.01968  -0.0113   1.0000   1.0000
  -1.000  -0.0721   0.02886   0.01940  -0.0086   1.0000   1.0000
  -0.750  -0.0714   0.02862   0.01912  -0.0059   1.0000   1.0000
  -0.500  -0.0704   0.02838   0.01884  -0.0032   1.0000   1.0000
  -0.250  -0.0691   0.02815   0.01858  -0.0006   1.0000   1.0000
   0.000  -0.0671   0.02794   0.01834   0.0018   1.0000   1.0000
   0.250  -0.0628   0.02778   0.01816   0.0039   1.0000   1.0000
   0.500  -0.0528   0.02779   0.01814   0.0048   1.0000   1.0000
   0.750  -0.0372   0.02799   0.01830   0.0048   1.0000   1.0000
   1.000  -0.0178   0.02835   0.01863   0.0040   1.0000   1.0000
   1.250   0.0042   0.02886   0.01912   0.0027   1.0000   1.0000
   1.500   0.0277   0.02949   0.01973   0.0011   1.0000   1.0000
   1.750   0.0520   0.03024   0.02048  -0.0007   1.0000   1.0000
   2.000   0.0765   0.03110   0.02136  -0.0027   1.0000   1.0000
   2.250   0.1146   0.03241   0.02269  -0.0073   0.9944   1.0000
   2.500   0.1897   0.03438   0.02470  -0.0179   0.9647   1.0000
   2.750   0.2579   0.03574   0.02613  -0.0262   0.9321   1.0000
   3.000   0.3230   0.03654   0.02703  -0.0330   0.8985   1.0000
   3.250   0.3848   0.03677   0.02738  -0.0383   0.8655   1.0000
   3.500   0.4452   0.03639   0.02715  -0.0423   0.8330   1.0000
   3.750   0.5053   0.03531   0.02626  -0.0452   0.8011   1.0000
   4.000   0.5661   0.03340   0.02457  -0.0469   0.7690   1.0000
   4.250   0.6239   0.03069   0.02209  -0.0470   0.7341   1.0000
   4.500   0.6770   0.02731   0.01893  -0.0452   0.6887   1.0000
   4.750   0.7216   0.02400   0.01550  -0.0416   0.5951   1.0000
   5.000   0.7555   0.02392   0.01413  -0.0390   0.4648   1.0000
   5.250   0.7822   0.02551   0.01498  -0.0385   0.4010   1.0000
   5.500   0.8133   0.02700   0.01616  -0.0392   0.3622   1.0000
   5.750   0.8472   0.02844   0.01729  -0.0404   0.3353   1.0000
   6.000   0.8786   0.02987   0.01867  -0.0413   0.3152   1.0000
   6.250   0.9120   0.03144   0.02015  -0.0425   0.3007   1.0000
   6.500   0.9419   0.03299   0.02178  -0.0433   0.2881   1.0000
   6.750   0.9692   0.03467   0.02362  -0.0438   0.2773   1.0000
   7.000   0.9997   0.03648   0.02538  -0.0447   0.2685   1.0000
   7.250   1.0233   0.03851   0.02781  -0.0447   0.2619   1.0000
   7.500   1.0511   0.04049   0.02984  -0.0453   0.2554   1.0000
   7.750   1.0714   0.04285   0.03252  -0.0450   0.2497   1.0000
   8.000   1.0892   0.04524   0.03526  -0.0445   0.2441   1.0000
   8.250   1.1110   0.04775   0.03796  -0.0445   0.2401   1.0000
   8.500   1.1329   0.05068   0.04100  -0.0446   0.2367   1.0000
   8.750   1.1358   0.05414   0.04503  -0.0428   0.2343   1.0000
   9.000   1.1344   0.05795   0.04930  -0.0411   0.2318   1.0000
   9.250   1.1292   0.06213   0.05387  -0.0393   0.2299   1.0000
   9.500   1.1136   0.06717   0.05928  -0.0374   0.2301   1.0000
   9.750   1.0849   0.07320   0.06561  -0.0355   0.2319   1.0000
  10.000   1.0505   0.07967   0.07225  -0.0342   0.2342   1.0000
  10.250   1.0286   0.08643   0.07911  -0.0346   0.2363   1.0000
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