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MH 60 10.08% (mh60-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: MH 60 10.08% (mh60-il)
Reynolds number: 50,000
Max Cl/Cd: 27.55 at α=7°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-mh60-il-50000.txt
Download as CSV file: xf-mh60-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 60  10.08%                                   
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4834   0.10057   0.09461   0.0149   1.0000   0.2722
  -8.000  -0.4850   0.09725   0.09138   0.0134   1.0000   0.2811
  -7.750  -0.4997   0.09547   0.08975   0.0111   1.0000   0.2926
  -7.500  -0.4756   0.08976   0.08402   0.0115   1.0000   0.2965
  -7.250  -0.4710   0.08598   0.08031   0.0102   1.0000   0.3010
  -7.000  -0.5492   0.06228   0.05594  -0.0280   1.0000   0.1091
  -6.750  -0.5327   0.05773   0.05145  -0.0281   1.0000   0.1066
  -6.500  -0.5231   0.05268   0.04617  -0.0293   1.0000   0.1033
  -6.250  -0.5192   0.04662   0.03903  -0.0305   1.0000   0.0954
  -6.000  -0.5046   0.04279   0.03479  -0.0300   1.0000   0.0954
  -5.750  -0.4898   0.03933   0.03086  -0.0289   1.0000   0.0956
  -5.500  -0.4755   0.03623   0.02740  -0.0272   1.0000   0.0959
  -5.250  -0.4618   0.03338   0.02440  -0.0253   1.0000   0.0981
  -5.000  -0.4480   0.03146   0.02232  -0.0232   1.0000   0.1052
  -4.750  -0.4333   0.02942   0.02002  -0.0211   1.0000   0.1131
  -4.500  -0.4161   0.02747   0.01783  -0.0190   1.0000   0.1248
  -4.250  -0.3996   0.02587   0.01629  -0.0171   1.0000   0.1465
  -4.000  -0.3827   0.02421   0.01476  -0.0151   1.0000   0.1837
  -3.750  -0.3671   0.02233   0.01343  -0.0128   1.0000   0.2706
  -3.500  -0.3574   0.02105   0.01288  -0.0094   1.0000   0.3994
  -3.250  -0.3475   0.02043   0.01267  -0.0057   1.0000   0.4970
  -3.000  -0.3358   0.01987   0.01246  -0.0020   1.0000   0.5836
  -2.750  -0.3236   0.01929   0.01223   0.0020   1.0000   0.6723
  -2.500  -0.3098   0.01872   0.01206   0.0065   1.0000   0.7817
  -2.250  -0.1308   0.01871   0.01125  -0.0195   1.0000   1.0000
  -2.000  -0.1454   0.01845   0.01073  -0.0156   1.0000   1.0000
  -1.750  -0.1420   0.01853   0.01050  -0.0134   1.0000   1.0000
  -1.500  -0.1331   0.01872   0.01038  -0.0118   1.0000   1.0000
  -1.250  -0.1216   0.01900   0.01038  -0.0104   1.0000   1.0000
  -1.000  -0.1086   0.01933   0.01047  -0.0094   1.0000   1.0000
  -0.750  -0.0848   0.01982   0.01073  -0.0104   0.9964   1.0000
  -0.500  -0.0281   0.02077   0.01137  -0.0173   0.9811   1.0000
  -0.250   0.0239   0.02163   0.01202  -0.0232   0.9648   1.0000
   0.000   0.0757   0.02247   0.01269  -0.0289   0.9480   1.0000
   0.250   0.1287   0.02329   0.01336  -0.0345   0.9311   1.0000
   0.500   0.1823   0.02405   0.01403  -0.0400   0.9141   1.0000
   0.750   0.2233   0.02473   0.01465  -0.0431   0.8955   1.0000
   1.000   0.2712   0.02536   0.01525  -0.0471   0.8772   1.0000
   1.250   0.3188   0.02591   0.01580  -0.0505   0.8596   1.0000
   1.500   0.3465   0.02658   0.01645  -0.0506   0.8402   1.0000
   1.750   0.3781   0.02715   0.01702  -0.0511   0.8212   1.0000
   2.000   0.4126   0.02758   0.01747  -0.0514   0.8034   1.0000
   2.250   0.4345   0.02822   0.01812  -0.0502   0.7838   1.0000
   2.500   0.4593   0.02874   0.01868  -0.0491   0.7647   1.0000
   2.750   0.4869   0.02906   0.01902  -0.0479   0.7471   1.0000
   3.000   0.5089   0.02955   0.01955  -0.0462   0.7282   1.0000
   3.250   0.5301   0.03004   0.02007  -0.0445   0.7087   1.0000
   3.500   0.5552   0.03021   0.02032  -0.0426   0.6909   1.0000
   3.750   0.5808   0.03022   0.02038  -0.0404   0.6742   1.0000
   4.000   0.5991   0.03086   0.02109  -0.0387   0.6529   1.0000
   4.250   0.6234   0.03088   0.02116  -0.0364   0.6348   1.0000
   4.500   0.6491   0.03064   0.02100  -0.0339   0.6179   1.0000
   4.750   0.6689   0.03115   0.02160  -0.0322   0.5964   1.0000
   5.000   0.6940   0.03092   0.02143  -0.0297   0.5778   1.0000
   5.250   0.7186   0.03075   0.02134  -0.0273   0.5588   1.0000
   5.500   0.7409   0.03096   0.02164  -0.0253   0.5371   1.0000
   5.750   0.7679   0.03043   0.02111  -0.0226   0.5183   1.0000
   6.000   0.7889   0.03102   0.02181  -0.0210   0.4947   1.0000
   6.250   0.8136   0.03100   0.02180  -0.0188   0.4732   1.0000
   6.500   0.8369   0.03136   0.02225  -0.0170   0.4503   1.0000
   6.750   0.8590   0.03197   0.02292  -0.0154   0.4268   1.0000
   7.000   0.8849   0.03212   0.02299  -0.0134   0.4050   1.0000
   7.250   0.9042   0.03331   0.02434  -0.0122   0.3810   1.0000
   7.500   0.9263   0.03419   0.02525  -0.0107   0.3590   1.0000
   7.750   0.9479   0.03519   0.02636  -0.0093   0.3369   1.0000
   8.000   0.9673   0.03652   0.02780  -0.0080   0.3156   1.0000
   8.250   0.9877   0.03780   0.02916  -0.0067   0.2951   1.0000
   8.500   1.0070   0.03918   0.03060  -0.0053   0.2750   1.0000
   8.750   1.0222   0.04113   0.03277  -0.0041   0.2562   1.0000
   9.000   1.0465   0.04193   0.03338  -0.0026   0.2354   1.0000
   9.250   1.0566   0.04455   0.03633  -0.0013   0.2193   1.0000
   9.500   1.0657   0.04696   0.03904   0.0001   0.2026   1.0000
   9.750   1.0744   0.04960   0.04188   0.0015   0.1879   1.0000
  10.000   1.0806   0.05241   0.04487   0.0029   0.1745   1.0000
  10.250   1.0852   0.05518   0.04780   0.0043   0.1620   1.0000
  10.500   1.0937   0.05820   0.05088   0.0055   0.1512   1.0000
  10.750   1.0705   0.06342   0.05650   0.0064   0.1498   1.0000
  11.000   1.0431   0.06874   0.06203   0.0067   0.1502   1.0000
  11.250   1.0139   0.07463   0.06801   0.0058   0.1514   1.0000
  11.500   0.9867   0.08154   0.07495   0.0034   0.1527   1.0000
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