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MH 23 8% (mh23-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: MH 23 8% (mh23-il)
Reynolds number: 100,000
Max Cl/Cd: 45.95 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-mh23-il-100000-n5.txt
Download as CSV file: xf-mh23-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4983   0.11030   0.10530  -0.0245   1.0000   0.0424
  -9.750  -0.4998   0.10655   0.10162  -0.0274   1.0000   0.0426
  -9.500  -0.5014   0.10256   0.09768  -0.0302   1.0000   0.0428
  -9.250  -0.5033   0.09848   0.09365  -0.0331   1.0000   0.0428
  -9.000  -0.5063   0.09426   0.08948  -0.0362   1.0000   0.0429
  -8.500  -0.4921   0.08480   0.07993  -0.0304   1.0000   0.0248
  -8.000  -0.4436   0.06426   0.05973  -0.0388   1.0000   0.0189
  -7.500  -0.5302   0.06772   0.06297  -0.0347   1.0000   0.0190
  -7.250  -0.5385   0.06441   0.05962  -0.0329   1.0000   0.0186
  -7.000  -0.5439   0.06126   0.05639  -0.0309   1.0000   0.0183
  -6.750  -0.5478   0.05800   0.05292  -0.0287   1.0000   0.0179
  -6.500  -0.5499   0.05459   0.04937  -0.0264   1.0000   0.0176
  -6.250  -0.5493   0.05126   0.04586  -0.0241   1.0000   0.0172
  -6.000  -0.5467   0.04788   0.04225  -0.0216   1.0000   0.0168
  -5.750  -0.5419   0.04451   0.03861  -0.0191   1.0000   0.0164
  -5.500  -0.5348   0.04123   0.03503  -0.0165   1.0000   0.0160
  -5.250  -0.5253   0.03814   0.03159  -0.0140   1.0000   0.0157
  -5.000  -0.5137   0.03512   0.02817  -0.0115   1.0000   0.0153
  -4.750  -0.4997   0.03232   0.02495  -0.0090   1.0000   0.0150
  -4.500  -0.4834   0.02971   0.02189  -0.0067   1.0000   0.0148
  -4.250  -0.4650   0.02739   0.01913  -0.0047   1.0000   0.0145
  -4.000  -0.4447   0.02535   0.01668  -0.0028   1.0000   0.0144
  -3.750  -0.4230   0.02350   0.01449  -0.0012   1.0000   0.0146
  -3.500  -0.4011   0.02201   0.01274   0.0003   1.0000   0.0147
  -3.250  -0.3794   0.02069   0.01124   0.0019   1.0000   0.0152
  -3.000  -0.3587   0.01957   0.00996   0.0035   1.0000   0.0159
  -2.750  -0.3375   0.01873   0.00895   0.0049   0.9997   0.0173
  -2.500  -0.3067   0.01791   0.00789   0.0041   0.9960   0.0230
  -2.250  -0.2754   0.01713   0.00701   0.0035   0.9922   0.0350
  -2.000  -0.2070   0.01351   0.00725  -0.0031   1.0000   0.9380
  -1.750  -0.1122   0.01395   0.00711  -0.0164   1.0000   1.0000
  -1.500  -0.0995   0.01394   0.00692  -0.0140   1.0000   1.0000
  -1.250  -0.0716   0.01400   0.00672  -0.0145   0.9967   1.0000
  -1.000  -0.0342   0.01411   0.00663  -0.0170   0.9910   1.0000
  -0.750   0.0037   0.01422   0.00658  -0.0195   0.9849   1.0000
  -0.500   0.0423   0.01434   0.00656  -0.0222   0.9788   1.0000
  -0.250   0.0807   0.01443   0.00651  -0.0248   0.9719   1.0000
   0.000   0.1188   0.01450   0.00651  -0.0272   0.9647   1.0000
   0.250   0.1595   0.01456   0.00653  -0.0301   0.9576   1.0000
   0.500   0.1964   0.01458   0.00652  -0.0322   0.9488   1.0000
   0.750   0.2430   0.01457   0.00653  -0.0362   0.9425   1.0000
   1.250   0.3153   0.01443   0.00644  -0.0396   0.9211   1.0000
   1.500   0.3503   0.01429   0.00637  -0.0409   0.9090   1.0000
   1.750   0.3823   0.01412   0.00630  -0.0415   0.8951   1.0000
   2.000   0.4115   0.01394   0.00619  -0.0414   0.8794   1.0000
   2.250   0.4365   0.01376   0.00609  -0.0405   0.8600   1.0000
   2.500   0.4613   0.01356   0.00597  -0.0394   0.8371   1.0000
   2.750   0.4867   0.01335   0.00585  -0.0383   0.8093   1.0000
   3.000   0.5170   0.01309   0.00573  -0.0380   0.7746   1.0000
   3.250   0.5501   0.01284   0.00546  -0.0380   0.7269   1.0000
   3.500   0.5791   0.01283   0.00530  -0.0374   0.6645   1.0000
   3.750   0.6024   0.01311   0.00536  -0.0358   0.5956   1.0000
   4.000   0.6223   0.01356   0.00556  -0.0339   0.5248   1.0000
   4.250   0.6406   0.01412   0.00593  -0.0320   0.4563   1.0000
   4.500   0.6582   0.01474   0.00631  -0.0300   0.3897   1.0000
   4.750   0.6754   0.01542   0.00675  -0.0281   0.3259   1.0000
   5.000   0.6925   0.01614   0.00726  -0.0262   0.2653   1.0000
   5.250   0.7096   0.01692   0.00783  -0.0245   0.2092   1.0000
   5.500   0.7266   0.01776   0.00847  -0.0227   0.1586   1.0000
   5.750   0.7431   0.01871   0.00920  -0.0210   0.1133   1.0000
   6.000   0.7596   0.01969   0.01010  -0.0191   0.0831   1.0000
   6.250   0.7757   0.02075   0.01106  -0.0172   0.0594   1.0000
   6.500   0.7909   0.02192   0.01228  -0.0150   0.0474   1.0000
   6.750   0.8056   0.02317   0.01377  -0.0127   0.0403   1.0000
   7.000   0.8198   0.02472   0.01549  -0.0102   0.0354   1.0000
   7.250   0.8348   0.02596   0.01680  -0.0083   0.0281   1.0000
   7.500   0.8507   0.02797   0.01906  -0.0060   0.0247   1.0000
   7.750   0.8661   0.02955   0.02084  -0.0040   0.0194   1.0000
   8.000   0.8782   0.03205   0.02358  -0.0018   0.0160   1.0000
   8.250   0.8920   0.03485   0.02681   0.0006   0.0145   1.0000
   8.500   0.9024   0.03741   0.02978   0.0031   0.0128   1.0000
   8.750   0.9104   0.03928   0.03196   0.0055   0.0110   1.0000
   9.000   0.9153   0.04117   0.03407   0.0080   0.0099   1.0000
   9.750   0.8916   0.05084   0.04467   0.0183   0.0086   1.0000
  10.000   0.8783   0.05379   0.04789   0.0215   0.0085   1.0000
  10.250   0.8624   0.05723   0.05156   0.0236   0.0085   1.0000
  10.500   0.8463   0.06101   0.05552   0.0245   0.0086   1.0000
  10.750   0.8288   0.06549   0.06019   0.0237   0.0086   1.0000
  11.000   0.8100   0.07115   0.06601   0.0208   0.0086   1.0000
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