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

Ornithopter airfoil. (s1020-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: Ornithopter airfoil. (s1020-il)
Reynolds number: 200,000
Max Cl/Cd: 81.49 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-s1020-il-200000-n5.txt
Download as CSV file: xf-s1020-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: Ornithopter airfoil.                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.1942   0.09034   0.08661  -0.0974   0.9557   0.0185
 -10.250  -0.1934   0.08421   0.08048  -0.1019   0.9514   0.0182
 -10.000  -0.2040   0.07657   0.07288  -0.1063   0.9439   0.0179
  -9.750  -0.2616   0.05568   0.05180  -0.1231   0.9332   0.0170
  -9.500  -0.2932   0.04599   0.04176  -0.1341   0.9212   0.0165
  -9.250  -0.3060   0.04075   0.03614  -0.1385   0.9100   0.0165
  -9.000  -0.3079   0.03695   0.03195  -0.1396   0.9003   0.0165
  -8.750  -0.2919   0.03343   0.02799  -0.1423   0.8957   0.0166
  -8.500  -0.2770   0.03084   0.02503  -0.1430   0.8892   0.0167
  -8.250  -0.2560   0.02858   0.02242  -0.1440   0.8837   0.0170
  -8.000  -0.2267   0.02652   0.02002  -0.1461   0.8804   0.0173
  -7.750  -0.2035   0.02510   0.01833  -0.1464   0.8748   0.0180
  -7.500  -0.1783   0.02380   0.01675  -0.1468   0.8693   0.0187
  -7.250  -0.1459   0.02251   0.01519  -0.1485   0.8657   0.0191
  -7.000  -0.1195   0.02132   0.01386  -0.1489   0.8606   0.0194
  -6.750  -0.0958   0.02023   0.01271  -0.1488   0.8545   0.0198
  -6.500  -0.0644   0.01924   0.01163  -0.1502   0.8505   0.0203
  -6.250  -0.0388   0.01851   0.01082  -0.1502   0.8449   0.0208
  -6.000  -0.0135   0.01783   0.01007  -0.1502   0.8388   0.0214
  -5.750   0.0182   0.01712   0.00924  -0.1515   0.8346   0.0224
  -5.500   0.0422   0.01660   0.00863  -0.1511   0.8283   0.0232
  -5.250   0.0693   0.01609   0.00801  -0.1514   0.8226   0.0245
  -5.000   0.1017   0.01556   0.00740  -0.1527   0.8185   0.0274
  -4.750   0.1244   0.01524   0.00700  -0.1519   0.8117   0.0298
  -4.500   0.1529   0.01479   0.00650  -0.1524   0.8063   0.0344
  -4.250   0.1830   0.01430   0.00601  -0.1532   0.8016   0.0500
  -4.000   0.2063   0.01385   0.00570  -0.1527   0.7950   0.0874
  -3.750   0.2349   0.01329   0.00537  -0.1535   0.7899   0.1555
  -3.500   0.2616   0.01273   0.00515  -0.1539   0.7845   0.2489
  -3.250   0.2866   0.01237   0.00506  -0.1538   0.7783   0.3315
  -3.000   0.3164   0.01217   0.00500  -0.1544   0.7734   0.4015
  -2.750   0.3419   0.01213   0.00500  -0.1540   0.7674   0.4423
  -2.500   0.3691   0.01211   0.00495  -0.1539   0.7615   0.4690
  -2.250   0.3997   0.01207   0.00485  -0.1545   0.7569   0.4896
  -2.000   0.4235   0.01209   0.00486  -0.1537   0.7501   0.5068
  -1.750   0.4516   0.01208   0.00481  -0.1538   0.7446   0.5234
  -1.500   0.4795   0.01209   0.00477  -0.1538   0.7392   0.5384
  -1.250   0.5047   0.01211   0.00477  -0.1533   0.7326   0.5524
  -1.000   0.5337   0.01211   0.00471  -0.1536   0.7274   0.5663
  -0.750   0.5588   0.01213   0.00474  -0.1530   0.7211   0.5784
  -0.500   0.5854   0.01215   0.00474  -0.1528   0.7150   0.5902
  -0.250   0.6136   0.01217   0.00471  -0.1529   0.7097   0.6026
   0.000   0.6379   0.01222   0.00477  -0.1523   0.7029   0.6153
   0.250   0.6656   0.01224   0.00477  -0.1523   0.6973   0.6276
   0.500   0.6906   0.01228   0.00483  -0.1518   0.6909   0.6401
   0.750   0.7163   0.01232   0.00488  -0.1514   0.6845   0.6531
   1.000   0.7434   0.01237   0.00491  -0.1513   0.6789   0.6667
   1.250   0.7672   0.01243   0.00501  -0.1505   0.6719   0.6808
   1.750   0.8175   0.01255   0.00517  -0.1496   0.6594   0.7107
   2.000   0.8424   0.01261   0.00526  -0.1490   0.6529   0.7269
   2.250   0.8668   0.01268   0.00536  -0.1483   0.6466   0.7441
   2.500   0.8897   0.01275   0.00548  -0.1474   0.6397   0.7623
   2.750   0.9144   0.01281   0.00556  -0.1467   0.6337   0.7822
   3.000   0.9349   0.01289   0.00573  -0.1453   0.6263   0.8049
   3.250   0.9582   0.01295   0.00581  -0.1443   0.6201   0.8300
   3.500   0.9770   0.01300   0.00596  -0.1424   0.6128   0.8619
   3.750   0.9980   0.01302   0.00604  -0.1409   0.6061   0.9104
   4.000   1.0260   0.01308   0.00614  -0.1412   0.5989   1.0000
   4.250   1.0507   0.01326   0.00629  -0.1408   0.5917   1.0000
   4.500   1.0746   0.01346   0.00647  -0.1403   0.5846   1.0000
   4.750   1.0979   0.01366   0.00667  -0.1396   0.5769   1.0000
   5.000   1.1209   0.01386   0.00687  -0.1389   0.5696   1.0000
   5.250   1.1423   0.01407   0.00707  -0.1378   0.5617   1.0000
   5.500   1.1636   0.01429   0.00731  -0.1367   0.5539   1.0000
   5.750   1.1841   0.01453   0.00754  -0.1355   0.5456   1.0000
   6.000   1.2040   0.01478   0.00781  -0.1342   0.5373   1.0000
   6.250   1.2245   0.01505   0.00807  -0.1330   0.5290   1.0000
   6.500   1.2432   0.01533   0.00840  -0.1315   0.5204   1.0000
   6.750   1.2631   0.01562   0.00868  -0.1302   0.5120   1.0000
   7.000   1.2804   0.01594   0.00905  -0.1285   0.5026   1.0000
   7.250   1.2987   0.01627   0.00940  -0.1269   0.4938   1.0000
   7.500   1.3154   0.01663   0.00979  -0.1252   0.4839   1.0000
   7.750   1.3314   0.01701   0.01021  -0.1233   0.4737   1.0000
   8.000   1.3470   0.01742   0.01064  -0.1214   0.4635   1.0000
   8.250   1.3613   0.01788   0.01111  -0.1193   0.4523   1.0000
   8.500   1.3749   0.01836   0.01166  -0.1171   0.4406   1.0000
   8.750   1.3878   0.01890   0.01222  -0.1149   0.4285   1.0000
   9.000   1.3988   0.01951   0.01285  -0.1124   0.4149   1.0000
   9.250   1.4088   0.02019   0.01354  -0.1098   0.4006   1.0000
   9.500   1.4167   0.02098   0.01433  -0.1071   0.3840   1.0000
   9.750   1.4219   0.02193   0.01526  -0.1041   0.3650   1.0000
  10.000   1.4235   0.02311   0.01637  -0.1007   0.3421   1.0000
  10.250   1.4235   0.02448   0.01764  -0.0974   0.3180   1.0000
  10.500   1.4217   0.02607   0.01913  -0.0941   0.2919   1.0000
  10.750   1.4196   0.02780   0.02076  -0.0910   0.2677   1.0000
  11.000   1.4148   0.02985   0.02268  -0.0879   0.2413   1.0000
  11.250   1.4151   0.03169   0.02447  -0.0855   0.2207   1.0000
  11.500   1.4128   0.03381   0.02651  -0.0831   0.1995   1.0000
  11.750   1.4116   0.03597   0.02862  -0.0810   0.1794   1.0000
  12.000   1.4102   0.03826   0.03085  -0.0790   0.1601   1.0000
  12.250   1.4070   0.04079   0.03330  -0.0772   0.1405   1.0000
  12.500   1.4044   0.04338   0.03583  -0.0756   0.1214   1.0000
  12.750   1.4004   0.04622   0.03859  -0.0741   0.1032   1.0000
  13.000   1.3989   0.04893   0.04126  -0.0728   0.0882   1.0000
  13.250   1.3972   0.05174   0.04406  -0.0718   0.0745   1.0000
  13.500   1.3950   0.05470   0.04699  -0.0709   0.0619   1.0000
  13.750   1.3924   0.05780   0.05007  -0.0701   0.0502   1.0000
  14.000   1.3904   0.06093   0.05320  -0.0695   0.0409   1.0000
  14.250   1.3876   0.06423   0.05653  -0.0690   0.0327   1.0000
  14.500   1.3850   0.06761   0.05995  -0.0687   0.0263   1.0000
  14.750   1.3828   0.07106   0.06344  -0.0685   0.0219   1.0000
  15.000   1.3799   0.07465   0.06710  -0.0684   0.0189   1.0000
  15.250   1.3792   0.07804   0.07059  -0.0685   0.0171   1.0000
  15.500   1.3772   0.08167   0.07432  -0.0687   0.0158   1.0000
  15.750   1.3748   0.08544   0.07818  -0.0691   0.0150   1.0000
  16.000   1.3734   0.08915   0.08203  -0.0695   0.0142   1.0000
  16.250   1.3718   0.09292   0.08593  -0.0702   0.0135   1.0000
  16.500   1.3694   0.09686   0.09000  -0.0709   0.0130   1.0000
  16.750   1.3661   0.10099   0.09425  -0.0719   0.0125   1.0000
  17.000   1.3618   0.10536   0.09872  -0.0730   0.0121   1.0000
<< Back to Ornithopter airfoil. (s1020-il)

Polar data table (+)

Polar graphs


<< Back to Ornithopter airfoil. (s1020-il)