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NLR-7301 AIRFOIL (nlr7301-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NLR-7301 AIRFOIL (nlr7301-il)
Reynolds number: 50,000
Max Cl/Cd: 20.98 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-nlr7301-il-50000.txt
Download as CSV file: xf-nlr7301-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NLR-7301 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.5529   0.08882   0.08060  -0.0504   1.0000   0.1912
 -10.000  -0.5397   0.08655   0.07838  -0.0488   1.0000   0.1930
  -9.750  -0.7665   0.07216   0.06377  -0.0470   1.0000   0.1743
  -9.500  -0.7876   0.06882   0.06037  -0.0440   1.0000   0.1736
  -9.250  -0.8089   0.06547   0.05688  -0.0410   1.0000   0.1729
  -9.000  -0.8269   0.06208   0.05329  -0.0382   1.0000   0.1724
  -8.750  -0.8401   0.05873   0.04967  -0.0358   1.0000   0.1723
  -8.500  -0.8476   0.05552   0.04613  -0.0337   1.0000   0.1730
  -8.250  -0.8505   0.05250   0.04266  -0.0320   1.0000   0.1743
  -8.000  -0.8386   0.04988   0.04005  -0.0308   1.0000   0.1771
  -7.750  -0.8250   0.04779   0.03792  -0.0295   1.0000   0.1803
  -7.500  -0.8125   0.04557   0.03548  -0.0283   1.0000   0.1837
  -7.250  -0.7992   0.04337   0.03292  -0.0275   1.0000   0.1881
  -7.000  -0.7825   0.04126   0.03062  -0.0267   1.0000   0.1929
  -6.750  -0.7629   0.03971   0.02912  -0.0256   1.0000   0.1989
  -6.500  -0.7436   0.03811   0.02723  -0.0249   1.0000   0.2063
  -6.250  -0.7225   0.03665   0.02591  -0.0238   1.0000   0.2144
  -6.000  -0.7013   0.03537   0.02445  -0.0231   1.0000   0.2248
  -5.750  -0.6803   0.03422   0.02357  -0.0216   1.0000   0.2363
  -5.500  -0.6594   0.03313   0.02260  -0.0202   1.0000   0.2508
  -5.250  -0.6393   0.03212   0.02174  -0.0186   1.0000   0.2691
  -5.000  -0.6201   0.03106   0.02097  -0.0170   1.0000   0.2936
  -4.750  -0.6019   0.02979   0.02015  -0.0155   1.0000   0.3303
  -4.500  -0.5867   0.02849   0.01992  -0.0131   1.0000   0.4057
  -4.250  -0.5810   0.03086   0.02335  -0.0039   1.0000   0.5223
  -4.000  -0.5714   0.03453   0.02702   0.0051   1.0000   0.5873
  -3.750  -0.5602   0.03739   0.02979   0.0126   1.0000   0.6261
  -3.500  -0.5478   0.03995   0.03228   0.0204   1.0000   0.6521
  -3.250  -0.5338   0.04179   0.03404   0.0268   1.0000   0.6781
  -3.000  -0.5188   0.04303   0.03519   0.0322   1.0000   0.7040
  -2.750  -0.5043   0.04377   0.03584   0.0368   1.0000   0.7297
  -2.500  -0.4917   0.04407   0.03605   0.0408   1.0000   0.7551
  -2.250  -0.4726   0.04437   0.03628   0.0447   1.0000   0.7790
  -2.000  -0.4430   0.04466   0.03648   0.0471   1.0000   0.8052
  -1.750  -0.4293   0.04441   0.03616   0.0501   1.0000   0.8326
  -1.500  -0.2923   0.04531   0.03686   0.0363   1.0000   0.8875
  -1.250  -0.2494   0.04469   0.03616   0.0328   1.0000   0.9085
  -1.000  -0.2189   0.04403   0.03545   0.0308   1.0000   0.9226
  -0.750  -0.1921   0.04343   0.03481   0.0292   1.0000   0.9337
  -0.500  -0.1786   0.04292   0.03427   0.0297   1.0000   0.9424
  -0.250  -0.1497   0.04249   0.03383   0.0274   1.0000   0.9508
   0.000  -0.1316   0.04212   0.03346   0.0270   1.0000   0.9574
   0.250  -0.1097   0.04186   0.03320   0.0258   1.0000   0.9636
   0.500  -0.0946   0.04165   0.03301   0.0257   1.0000   0.9690
   0.750  -0.0710   0.04156   0.03295   0.0240   1.0000   0.9740
   1.000  -0.0553   0.04154   0.03297   0.0237   1.0000   0.9788
   1.250  -0.0346   0.04162   0.03310   0.0224   1.0000   0.9832
   1.500  -0.0150   0.04181   0.03335   0.0212   1.0000   0.9874
   1.750   0.0016   0.04209   0.03370   0.0204   1.0000   0.9918
   2.000   0.0212   0.04250   0.03420   0.0188   1.0000   0.9958
   2.250   0.4038   0.04045   0.03270  -0.0376   0.8687   1.0000
   2.500   0.5755   0.03097   0.02381  -0.0487   0.7560   1.0000
   2.750   0.5958   0.02932   0.01940  -0.0385   0.3604   1.0000
   3.000   0.6070   0.03069   0.02009  -0.0360   0.3114   1.0000
   3.250   0.6273   0.03155   0.02069  -0.0349   0.2803   1.0000
   3.500   0.6534   0.03231   0.02124  -0.0348   0.2579   1.0000
   3.750   0.6808   0.03310   0.02185  -0.0348   0.2409   1.0000
   4.000   0.7096   0.03405   0.02252  -0.0352   0.2278   1.0000
   4.250   0.7309   0.03485   0.02344  -0.0343   0.2177   1.0000
   4.500   0.7568   0.03608   0.02445  -0.0344   0.2097   1.0000
   4.750   0.7742   0.03702   0.02566  -0.0329   0.2034   1.0000
   5.000   0.7933   0.03810   0.02678  -0.0319   0.1975   1.0000
   5.250   0.8144   0.03961   0.02820  -0.0314   0.1930   1.0000
   5.500   0.8287   0.04107   0.02992  -0.0296   0.1905   1.0000
   5.750   0.8403   0.04254   0.03170  -0.0274   0.1882   1.0000
   6.000   0.8498   0.04409   0.03352  -0.0251   0.1859   1.0000
   6.250   0.8573   0.04568   0.03533  -0.0225   0.1837   1.0000
   6.500   0.8633   0.04730   0.03715  -0.0199   0.1818   1.0000
   6.750   0.8676   0.04900   0.03902  -0.0170   0.1804   1.0000
   7.000   0.8683   0.05083   0.04106  -0.0137   0.1797   1.0000
   7.250   0.8620   0.05282   0.04334  -0.0096   0.1801   1.0000
   7.500   0.8493   0.05489   0.04571  -0.0049   0.1810   1.0000
   7.750   0.8313   0.05697   0.04808   0.0003   0.1823   1.0000
   8.000   0.8090   0.05906   0.05042   0.0057   0.1839   1.0000
   8.250   0.7885   0.06175   0.05338   0.0098   0.1859   1.0000
   8.500   0.7766   0.06540   0.05727   0.0116   0.1882   1.0000
   8.750   0.7746   0.06956   0.06158   0.0117   0.1905   1.0000
   9.000   0.7910   0.07403   0.06607   0.0104   0.1923   1.0000
   9.250   0.6638   0.08692   0.07969   0.0076   0.2117   1.0000
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