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DSMA-523A AIRFOIL (dsma523a-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: DSMA-523A AIRFOIL (dsma523a-il)
Reynolds number: 100,000
Max Cl/Cd: 20.63 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-dsma523a-il-100000.txt
Download as CSV file: xf-dsma523a-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: DSMA-523A 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5525   0.09745   0.09177  -0.0173   1.0000   0.1934
  -8.500  -0.7998   0.05822   0.05153  -0.0435   1.0000   0.1009
  -8.250  -0.7920   0.05410   0.04732  -0.0433   1.0000   0.0995
  -8.000  -0.7853   0.04948   0.04240  -0.0438   1.0000   0.0975
  -7.750  -0.7752   0.04388   0.03620  -0.0452   1.0000   0.0944
  -7.500  -0.7551   0.03880   0.03002  -0.0475   1.0000   0.0914
  -7.250  -0.7323   0.03594   0.02687  -0.0478   1.0000   0.0915
  -7.000  -0.7094   0.03362   0.02443  -0.0477   1.0000   0.0927
  -6.750  -0.6860   0.03216   0.02298  -0.0473   1.0000   0.0952
  -6.500  -0.6608   0.03056   0.02119  -0.0474   1.0000   0.0980
  -6.250  -0.6344   0.02893   0.01928  -0.0473   1.0000   0.1001
  -6.000  -0.6082   0.02745   0.01755  -0.0471   1.0000   0.1024
  -5.750  -0.5848   0.02617   0.01642  -0.0464   1.0000   0.1064
  -5.500  -0.5591   0.02522   0.01542  -0.0460   1.0000   0.1113
  -5.250  -0.5335   0.02422   0.01431  -0.0453   1.0000   0.1154
  -5.000  -0.5098   0.02325   0.01352  -0.0445   1.0000   0.1214
  -4.750  -0.4836   0.02258   0.01279  -0.0441   1.0000   0.1287
  -4.500  -0.4598   0.02168   0.01212  -0.0433   1.0000   0.1364
  -4.250  -0.4339   0.02105   0.01154  -0.0428   1.0000   0.1472
  -4.000  -0.4075   0.02040   0.01106  -0.0426   1.0000   0.1614
  -3.750  -0.3802   0.01967   0.01063  -0.0426   1.0000   0.1857
  -3.500  -0.3356   0.01738   0.01039  -0.0477   1.0000   0.4450
  -3.250  -0.3594   0.01955   0.01331  -0.0321   1.0000   0.5366
  -3.000  -0.3271   0.02090   0.01449  -0.0323   1.0000   0.6419
  -2.750  -0.3027   0.02197   0.01542  -0.0307   1.0000   0.6725
  -2.500  -0.2896   0.02291   0.01632  -0.0261   1.0000   0.6903
  -2.250  -0.2819   0.02358   0.01699  -0.0202   1.0000   0.7013
  -2.000  -0.2694   0.02408   0.01747  -0.0158   1.0000   0.7153
  -1.750  -0.2548   0.02445   0.01781  -0.0121   1.0000   0.7305
  -1.500  -0.2394   0.02475   0.01809  -0.0087   1.0000   0.7463
  -1.250  -0.2261   0.02492   0.01825  -0.0049   1.0000   0.7600
  -1.000  -0.2157   0.02488   0.01821  -0.0005   1.0000   0.7714
  -0.750  -0.1997   0.02486   0.01818   0.0023   1.0000   0.7852
  -0.500  -0.1850   0.02483   0.01816   0.0054   1.0000   0.8007
  -0.250  -0.1775   0.02469   0.01805   0.0104   1.0000   0.8186
   0.000  -0.1727   0.02440   0.01779   0.0160   1.0000   0.8377
   0.250  -0.1660   0.02402   0.01745   0.0209   1.0000   0.8567
   0.500  -0.1530   0.02366   0.01710   0.0239   1.0000   0.8713
   0.750  -0.1335   0.02341   0.01688   0.0250   1.0000   0.8816
   1.000  -0.1125   0.02321   0.01669   0.0257   1.0000   0.8893
   1.250  -0.0887   0.02306   0.01659   0.0256   1.0000   0.8943
   1.500  -0.0665   0.02294   0.01651   0.0259   1.0000   0.9000
   1.750  -0.0455   0.02282   0.01645   0.0264   1.0000   0.9068
   2.000  -0.0221   0.02287   0.01657   0.0262   1.0000   0.9139
   2.250   0.0532   0.02317   0.01698   0.0179   0.9708   0.9242
   2.500   0.1268   0.02234   0.01628   0.0107   0.9355   0.9274
   2.750   0.1926   0.02117   0.01526   0.0054   0.9023   0.9281
   3.000   0.2356   0.01973   0.01399   0.0047   0.8547   0.9286
   3.250   0.3513   0.02102   0.01197  -0.0061   0.2014   0.9301
   3.500   0.3726   0.02167   0.01239  -0.0053   0.1741   0.9308
   3.750   0.3968   0.02232   0.01291  -0.0050   0.1581   0.9314
   4.000   0.4247   0.02316   0.01356  -0.0054   0.1460   0.9314
   4.250   0.4546   0.02385   0.01425  -0.0061   0.1374   0.9317
   4.500   0.4874   0.02508   0.01529  -0.0075   0.1296   0.9324
   4.750   0.5176   0.02577   0.01611  -0.0080   0.1236   0.9326
   5.000   0.5498   0.02682   0.01721  -0.0091   0.1190   0.9324
   5.250   0.5830   0.02843   0.01872  -0.0107   0.1142   0.9323
   5.500   0.6121   0.02967   0.02021  -0.0111   0.1106   0.9322
   5.750   0.6408   0.03109   0.02192  -0.0115   0.1081   0.9321
   6.000   0.6675   0.03265   0.02375  -0.0114   0.1060   0.9326
   6.250   0.6928   0.03413   0.02543  -0.0114   0.1032   0.9333
   6.500   0.7193   0.03582   0.02716  -0.0118   0.1004   0.9335
   6.750   0.7433   0.03821   0.02985  -0.0117   0.0998   0.9334
   7.000   0.7604   0.04118   0.03349  -0.0103   0.1015   0.9333
   7.250   0.7702   0.04587   0.03896  -0.0083   0.1064   0.9331
   7.500   0.7827   0.04999   0.04346  -0.0074   0.1092   0.9331
   7.750   0.7982   0.05437   0.04803  -0.0074   0.1117   0.9331
   8.000   0.7130   0.07658   0.07221  -0.0103   0.2033   0.9326
   8.250   0.7011   0.08029   0.07604  -0.0121   0.1943   0.9326
   8.500   0.7448   0.08295   0.07861  -0.0108   0.1894   0.9332
   8.750   0.6919   0.09017   0.08597  -0.0170   0.1829   0.9334
   9.000   0.6774   0.09616   0.09197  -0.0241   0.1748   0.9337
   9.250   0.7047   0.09802   0.09385  -0.0212   0.1704   0.9341
   9.500   0.7576   0.10164   0.09737  -0.0150   0.1669   0.9345
   9.750   0.6742   0.11102   0.10681  -0.0348   0.1585   0.9337
  10.000   0.6850   0.11455   0.11034  -0.0360   0.1535   0.9338
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