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MH 23 8% (mh23-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: MH 23 8% (mh23-il)
Reynolds number: 50,000
Max Cl/Cd: 33.16 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-mh23-il-50000-n5.txt
Download as CSV file: xf-mh23-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 23  8%                                       
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4917   0.09597   0.08895  -0.0268   1.0000   0.0352
  -8.750  -0.4955   0.09152   0.08458  -0.0293   1.0000   0.0337
  -8.500  -0.5024   0.08670   0.07986  -0.0329   1.0000   0.0324
  -8.000  -0.5198   0.07842   0.07173  -0.0365   1.0000   0.0300
  -7.750  -0.5241   0.07502   0.06834  -0.0358   1.0000   0.0296
  -7.500  -0.5284   0.07172   0.06503  -0.0348   1.0000   0.0292
  -7.250  -0.5332   0.06830   0.06143  -0.0336   1.0000   0.0288
  -7.000  -0.5365   0.06503   0.05809  -0.0322   1.0000   0.0284
  -6.750  -0.5387   0.06169   0.05463  -0.0306   1.0000   0.0280
  -6.500  -0.5396   0.05826   0.05104  -0.0287   1.0000   0.0276
  -6.250  -0.5382   0.05494   0.04753  -0.0267   1.0000   0.0273
  -6.000  -0.5348   0.05162   0.04395  -0.0246   1.0000   0.0269
  -5.750  -0.5292   0.04837   0.04039  -0.0223   1.0000   0.0269
  -5.500  -0.5211   0.04524   0.03689  -0.0200   1.0000   0.0270
  -5.250  -0.5106   0.04230   0.03356  -0.0177   1.0000   0.0274
  -5.000  -0.4977   0.03949   0.03028  -0.0155   1.0000   0.0281
  -4.750  -0.4822   0.03694   0.02721  -0.0133   1.0000   0.0293
  -4.500  -0.4643   0.03471   0.02443  -0.0111   1.0000   0.0303
  -4.250  -0.4458   0.03208   0.02145  -0.0094   1.0000   0.0313
  -4.000  -0.4253   0.02985   0.01894  -0.0079   1.0000   0.0324
  -3.750  -0.4029   0.02789   0.01672  -0.0065   1.0000   0.0342
  -3.500  -0.3797   0.02628   0.01487  -0.0050   1.0000   0.0361
  -3.250  -0.3557   0.02485   0.01314  -0.0036   1.0000   0.0403
  -3.000  -0.3322   0.02363   0.01177  -0.0026   1.0000   0.0531
  -2.750  -0.3080   0.02224   0.01033  -0.0017   1.0000   0.0741
  -2.500  -0.1421   0.01814   0.00911  -0.0250   1.0000   1.0000
  -2.250  -0.1302   0.01803   0.00868  -0.0225   1.0000   1.0000
  -2.000  -0.1175   0.01795   0.00831  -0.0201   1.0000   1.0000
  -1.750  -0.1042   0.01790   0.00800  -0.0178   1.0000   1.0000
  -1.500  -0.0905   0.01788   0.00774  -0.0156   1.0000   1.0000
  -1.250  -0.0762   0.01789   0.00746  -0.0134   1.0000   1.0000
  -1.000  -0.0616   0.01793   0.00729  -0.0113   1.0000   1.0000
  -0.750  -0.0468   0.01799   0.00718  -0.0093   1.0000   1.0000
  -0.500  -0.0317   0.01807   0.00711  -0.0073   1.0000   1.0000
  -0.250  -0.0164   0.01818   0.00708  -0.0054   1.0000   1.0000
   0.000  -0.0010   0.01831   0.00710  -0.0036   1.0000   1.0000
   0.250   0.0144   0.01848   0.00711  -0.0018   1.0000   1.0000
   0.500   0.0299   0.01867   0.00722   0.0000   1.0000   1.0000
   0.750   0.0454   0.01890   0.00738   0.0016   1.0000   1.0000
   1.000   0.0785   0.01931   0.00775  -0.0003   0.9938   1.0000
   1.250   0.1231   0.01979   0.00823  -0.0045   0.9825   1.0000
   1.500   0.1675   0.02023   0.00871  -0.0086   0.9704   1.0000
   1.750   0.2119   0.02062   0.00915  -0.0125   0.9572   1.0000
   2.000   0.2567   0.02093   0.00957  -0.0163   0.9432   1.0000
   2.250   0.2990   0.02114   0.00990  -0.0194   0.9274   1.0000
   2.500   0.3388   0.02125   0.01022  -0.0219   0.9094   1.0000
   2.750   0.3831   0.02124   0.01041  -0.0249   0.8912   1.0000
   3.000   0.4189   0.02114   0.01051  -0.0262   0.8690   1.0000
   3.250   0.4543   0.02091   0.01052  -0.0270   0.8448   1.0000
   3.500   0.4865   0.02059   0.01056  -0.0269   0.8172   1.0000
   3.750   0.5201   0.02016   0.01040  -0.0268   0.7861   1.0000
   4.000   0.5543   0.01975   0.01022  -0.0266   0.7470   1.0000
   4.250   0.5906   0.01935   0.00999  -0.0264   0.6952   1.0000
   4.500   0.6240   0.01921   0.00995  -0.0256   0.6266   1.0000
   4.750   0.6479   0.01954   0.01009  -0.0236   0.5493   1.0000
   5.000   0.6658   0.02020   0.01049  -0.0210   0.4713   1.0000
   5.250   0.6802   0.02107   0.01106  -0.0182   0.3956   1.0000
   5.500   0.6936   0.02208   0.01179  -0.0156   0.3233   1.0000
   5.750   0.7069   0.02324   0.01267  -0.0132   0.2555   1.0000
   6.000   0.7210   0.02456   0.01371  -0.0111   0.1959   1.0000
   6.250   0.7365   0.02606   0.01500  -0.0093   0.1467   1.0000
   6.500   0.7529   0.02763   0.01643  -0.0077   0.1096   1.0000
   6.750   0.7700   0.02941   0.01814  -0.0061   0.0888   1.0000
   7.000   0.7893   0.03122   0.02009  -0.0045   0.0724   1.0000
   7.250   0.8072   0.03298   0.02219  -0.0029   0.0596   1.0000
   7.500   0.8279   0.03528   0.02470  -0.0015   0.0529   1.0000
   7.750   0.8499   0.03840   0.02816  -0.0002   0.0486   1.0000
   8.000   0.8653   0.04110   0.03133   0.0018   0.0429   1.0000
   8.250   0.8745   0.04359   0.03400   0.0038   0.0374   1.0000
   8.500   0.8827   0.04724   0.03826   0.0063   0.0353   1.0000
   8.750   0.8854   0.05107   0.04262   0.0091   0.0340   1.0000
   9.000   0.8833   0.05488   0.04685   0.0119   0.0333   1.0000
   9.250   0.8760   0.05875   0.05109   0.0146   0.0327   1.0000
   9.500   0.8649   0.06254   0.05517   0.0171   0.0325   1.0000
   9.750   0.8487   0.06614   0.05899   0.0197   0.0326   1.0000
  10.000   0.8303   0.06995   0.06296   0.0213   0.0328   1.0000
  10.250   0.8105   0.07439   0.06754   0.0212   0.0331   1.0000
  10.500   0.7923   0.07939   0.07263   0.0195   0.0337   1.0000
  10.750   0.7736   0.08559   0.07888   0.0158   0.0342   1.0000
  11.000   0.7574   0.09289   0.08620   0.0109   0.0348   1.0000
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