Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

MH 112 Airfoil (mh112-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: MH 112 Airfoil (mh112-il)
Reynolds number: 1,000,000
Max Cl/Cd: 126.59 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-mh112-il-1000000-n5.txt
Download as CSV file: xf-mh112-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 112  Airfoil                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.000  -0.0268   0.07373   0.07063  -0.1627   0.8354   0.0091
 -12.750  -0.0742   0.06331   0.06004  -0.1679   0.8259   0.0090
 -12.500  -0.0988   0.05703   0.05363  -0.1711   0.8177   0.0089
 -12.000  -0.1510   0.04496   0.04132  -0.1772   0.8004   0.0089
 -11.750  -0.1900   0.03759   0.03382  -0.1811   0.7920   0.0088
 -11.500  -0.2073   0.03286   0.02896  -0.1835   0.7839   0.0088
 -11.250  -0.2395   0.02702   0.02295  -0.1868   0.7763   0.0088
 -11.000  -0.2408   0.02327   0.01900  -0.1920   0.7689   0.0088
 -10.500  -0.2131   0.02042   0.01587  -0.1934   0.7588   0.0088
 -10.250  -0.1936   0.01944   0.01477  -0.1937   0.7540   0.0089
 -10.000  -0.1743   0.01836   0.01356  -0.1939   0.7490   0.0089
  -9.750  -0.1536   0.01739   0.01244  -0.1941   0.7441   0.0089
  -9.500  -0.1306   0.01648   0.01143  -0.1944   0.7404   0.0090
  -9.250  -0.1066   0.01570   0.01055  -0.1946   0.7359   0.0091
  -9.000  -0.0820   0.01505   0.00980  -0.1947   0.7312   0.0092
  -8.750  -0.0568   0.01449   0.00914  -0.1948   0.7267   0.0093
  -8.500  -0.0304   0.01399   0.00858  -0.1950   0.7228   0.0094
  -8.250  -0.0036   0.01355   0.00808  -0.1952   0.7185   0.0095
  -8.000   0.0233   0.01314   0.00761  -0.1953   0.7141   0.0096
  -7.750   0.0501   0.01278   0.00717  -0.1953   0.7094   0.0097
  -7.500   0.0773   0.01242   0.00675  -0.1954   0.7053   0.0098
  -7.250   0.1052   0.01206   0.00634  -0.1956   0.7013   0.0099
  -7.000   0.1329   0.01175   0.00597  -0.1958   0.6965   0.0100
  -6.750   0.1604   0.01147   0.00563  -0.1958   0.6918   0.0101
  -6.500   0.1881   0.01121   0.00531  -0.1959   0.6876   0.0103
  -6.250   0.2166   0.01094   0.00500  -0.1961   0.6836   0.0105
  -6.000   0.2448   0.01071   0.00472  -0.1962   0.6787   0.0106
  -5.750   0.2726   0.01050   0.00446  -0.1962   0.6738   0.0109
  -5.500   0.3005   0.01032   0.00422  -0.1962   0.6694   0.0111
  -5.250   0.3292   0.01014   0.00401  -0.1964   0.6654   0.0113
  -5.000   0.3576   0.00993   0.00377  -0.1965   0.6604   0.0117
  -4.750   0.3855   0.00978   0.00357  -0.1965   0.6550   0.0123
  -4.500   0.4135   0.00965   0.00341  -0.1965   0.6506   0.0130
  -4.250   0.4422   0.00952   0.00326  -0.1967   0.6463   0.0137
  -4.000   0.4705   0.00940   0.00312  -0.1967   0.6411   0.0147
  -3.750   0.4981   0.00932   0.00300  -0.1966   0.6356   0.0161
  -3.500   0.5263   0.00922   0.00288  -0.1967   0.6309   0.0181
  -3.250   0.5548   0.00912   0.00279  -0.1968   0.6260   0.0208
  -3.000   0.5827   0.00904   0.00270  -0.1967   0.6206   0.0254
  -2.750   0.6099   0.00899   0.00263  -0.1966   0.6151   0.0314
  -2.500   0.6384   0.00891   0.00257  -0.1967   0.6101   0.0394
  -2.250   0.6663   0.00884   0.00252  -0.1967   0.6047   0.0486
  -2.000   0.6934   0.00880   0.00248  -0.1966   0.5991   0.0609
  -1.750   0.7214   0.00873   0.00245  -0.1966   0.5940   0.0779
  -1.500   0.7493   0.00865   0.00244  -0.1967   0.5883   0.0996
  -1.250   0.7763   0.00861   0.00244  -0.1966   0.5824   0.1246
  -1.000   0.8036   0.00857   0.00245  -0.1965   0.5768   0.1495
  -0.750   0.8308   0.00854   0.00246  -0.1964   0.5698   0.1744
  -0.250   0.8833   0.00859   0.00254  -0.1958   0.5544   0.2191
   0.000   0.9088   0.00864   0.00260  -0.1954   0.5463   0.2426
   0.250   0.9352   0.00865   0.00266  -0.1951   0.5401   0.2719
   0.500   0.9615   0.00868   0.00272  -0.1948   0.5336   0.2992
   0.750   0.9863   0.00875   0.00280  -0.1943   0.5268   0.3250
   1.000   1.0125   0.00878   0.00288  -0.1940   0.5212   0.3537
   1.250   1.0381   0.00883   0.00297  -0.1936   0.5149   0.3823
   1.500   1.0612   0.00893   0.00308  -0.1927   0.5084   0.4084
   1.750   1.0859   0.00898   0.00317  -0.1921   0.5029   0.4354
   2.000   1.1094   0.00906   0.00328  -0.1913   0.4965   0.4604
   2.250   1.1311   0.00918   0.00341  -0.1901   0.4903   0.4827
   2.500   1.1552   0.00926   0.00354  -0.1894   0.4846   0.5082
   2.750   1.1779   0.00940   0.00368  -0.1885   0.4779   0.5295
   3.000   1.1996   0.00956   0.00385  -0.1873   0.4717   0.5497
   3.250   1.2233   0.00968   0.00400  -0.1866   0.4656   0.5719
   3.500   1.2445   0.00986   0.00420  -0.1854   0.4582   0.5929
   3.750   1.2662   0.01003   0.00438  -0.1843   0.4523   0.6146
   4.000   1.2887   0.01018   0.00457  -0.1834   0.4457   0.6386
   4.250   1.3082   0.01041   0.00482  -0.1819   0.4379   0.6653
   4.500   1.3301   0.01057   0.00503  -0.1809   0.4315   0.6949
   5.000   1.3693   0.01096   0.00557  -0.1782   0.4162   0.7839
   5.250   1.3812   0.01096   0.00585  -0.1751   0.4085   0.9275
   5.500   1.4005   0.01125   0.00612  -0.1737   0.4006   1.0000
   5.750   1.4194   0.01159   0.00642  -0.1723   0.3913   1.0000
   6.000   1.4363   0.01200   0.00678  -0.1706   0.3820   1.0000
   6.250   1.4538   0.01241   0.00715  -0.1691   0.3725   1.0000
   6.500   1.4704   0.01285   0.00755  -0.1674   0.3638   1.0000
   6.750   1.4864   0.01334   0.00799  -0.1657   0.3541   1.0000
   7.000   1.5018   0.01387   0.00847  -0.1639   0.3441   1.0000
   7.250   1.5159   0.01447   0.00903  -0.1620   0.3345   1.0000
   7.500   1.5318   0.01501   0.00954  -0.1604   0.3258   1.0000
   7.750   1.5454   0.01567   0.01015  -0.1585   0.3172   1.0000
   8.000   1.5607   0.01627   0.01073  -0.1569   0.3089   1.0000
   8.250   1.5735   0.01701   0.01144  -0.1550   0.3009   1.0000
   8.500   1.5893   0.01762   0.01204  -0.1535   0.2941   1.0000
   8.750   1.6014   0.01844   0.01283  -0.1516   0.2866   1.0000
   9.000   1.6168   0.01911   0.01349  -0.1503   0.2806   1.0000
   9.250   1.6293   0.01995   0.01432  -0.1485   0.2734   1.0000
   9.500   1.6423   0.02079   0.01514  -0.1469   0.2672   1.0000
   9.750   1.6564   0.02158   0.01594  -0.1455   0.2616   1.0000
  10.000   1.6679   0.02256   0.01690  -0.1438   0.2553   1.0000
  10.250   1.6810   0.02346   0.01780  -0.1424   0.2496   1.0000
  10.500   1.6934   0.02443   0.01877  -0.1410   0.2438   1.0000
  10.750   1.7031   0.02559   0.01992  -0.1393   0.2369   1.0000
  11.000   1.7162   0.02656   0.02091  -0.1380   0.2324   1.0000
  11.250   1.7262   0.02777   0.02210  -0.1365   0.2253   1.0000
  11.500   1.7365   0.02900   0.02333  -0.1350   0.2204   1.0000
  11.750   1.7479   0.03016   0.02450  -0.1338   0.2145   1.0000
  12.000   1.7549   0.03168   0.02601  -0.1322   0.2081   1.0000
  12.250   1.7663   0.03290   0.02725  -0.1310   0.2032   1.0000
  12.500   1.7746   0.03441   0.02875  -0.1297   0.1972   1.0000
  12.750   1.7823   0.03599   0.03034  -0.1283   0.1913   1.0000
  13.000   1.7909   0.03752   0.03188  -0.1271   0.1856   1.0000
  13.250   1.7963   0.03937   0.03373  -0.1258   0.1796   1.0000
  13.500   1.8057   0.04091   0.03530  -0.1248   0.1752   1.0000
  13.750   1.8114   0.04282   0.03721  -0.1236   0.1692   1.0000
  14.000   1.8174   0.04473   0.03913  -0.1225   0.1641   1.0000
  14.250   1.8237   0.04665   0.04107  -0.1215   0.1587   1.0000
  14.500   1.8266   0.04895   0.04337  -0.1204   0.1529   1.0000
  14.750   1.8335   0.05090   0.04534  -0.1196   0.1486   1.0000
  15.000   1.8375   0.05316   0.04762  -0.1188   0.1436   1.0000
  15.250   1.8414   0.05549   0.04997  -0.1180   0.1392   1.0000
  15.500   1.8465   0.05772   0.05223  -0.1173   0.1348   1.0000
  15.750   1.8483   0.06036   0.05489  -0.1166   0.1301   1.0000
  16.000   1.8528   0.06274   0.05731  -0.1160   0.1267   1.0000
  16.250   1.8561   0.06529   0.05989  -0.1156   0.1224   1.0000
  16.500   1.8562   0.06826   0.06288  -0.1151   0.1182   1.0000
  16.750   1.8602   0.07079   0.06546  -0.1147   0.1147   1.0000
  17.000   1.8612   0.07373   0.06843  -0.1144   0.1108   1.0000
  17.500   1.8635   0.07970   0.07448  -0.1141   0.1042   1.0000
  17.750   1.8646   0.08274   0.07755  -0.1140   0.1005   1.0000
  18.000   1.8621   0.08629   0.08114  -0.1141   0.0972   1.0000
  18.250   1.8645   0.08917   0.08408  -0.1142   0.0942   1.0000
  18.500   1.8628   0.09271   0.08766  -0.1144   0.0911   1.0000
  18.750   1.8608   0.09627   0.09126  -0.1147   0.0879   1.0000
  19.000   1.8611   0.09956   0.09461  -0.1150   0.0855   1.0000
  19.250   1.8597   0.10307   0.09816  -0.1155   0.0824   1.0000
<< Back to MH 112 Airfoil (mh112-il)

Polar data table (+)

Polar graphs


<< Back to MH 112 Airfoil (mh112-il)