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

GOE 363 AIRFOIL (goe363-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 363 AIRFOIL (goe363-il)
Reynolds number: 100,000
Max Cl/Cd: 53.33 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe363-il-100000.txt
Download as CSV file: xf-goe363-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 363 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.000  -0.3499   0.10237   0.09853  -0.0116   1.0000   0.0946
  -6.750  -0.3670   0.10148   0.09772  -0.0113   1.0000   0.0964
  -6.500  -0.3841   0.10082   0.09711  -0.0180   1.0000   0.0980
  -6.250  -0.3669   0.09520   0.09153  -0.0190   0.9963   0.0997
  -6.000  -0.3440   0.09133   0.08764  -0.0183   0.9922   0.1030
  -5.750  -0.2993   0.08649   0.08258  -0.0433   0.9805   0.1124
  -5.500  -0.2867   0.08174   0.07794  -0.0384   0.9759   0.1140
  -5.250  -0.2597   0.07826   0.07445  -0.0393   0.9710   0.1194
  -5.000  -0.2250   0.07331   0.06935  -0.0521   0.9607   0.1285
  -4.750  -0.1956   0.06971   0.06575  -0.0534   0.9560   0.1332
  -4.500  -0.1577   0.06519   0.06099  -0.0646   0.9453   0.1434
  -4.250  -0.1246   0.06154   0.05735  -0.0665   0.9409   0.1482
  -4.000  -0.0918   0.05785   0.05347  -0.0728   0.9301   0.1599
  -3.750  -0.0451   0.05429   0.04962  -0.0805   0.9233   0.1740
  -3.500   0.0150   0.04312   0.03735  -0.0919   0.9167   0.1129
  -3.250   0.0679   0.03649   0.02984  -0.0980   0.9128   0.0980
  -3.000   0.1062   0.03349   0.02649  -0.1006   0.9051   0.0965
  -2.750   0.1529   0.03089   0.02322  -0.1039   0.8986   0.0981
  -2.500   0.1943   0.02909   0.02125  -0.1063   0.8893   0.1017
  -2.250   0.2524   0.02701   0.01885  -0.1106   0.8816   0.1068
  -2.000   0.2988   0.02542   0.01705  -0.1127   0.8692   0.1160
  -1.750   0.3426   0.02397   0.01549  -0.1141   0.8584   0.1295
  -1.500   0.3864   0.02248   0.01406  -0.1156   0.8510   0.1616
  -1.250   0.4164   0.02168   0.01340  -0.1154   0.8394   0.2218
  -1.000   0.4513   0.02104   0.01281  -0.1157   0.8304   0.2568
  -0.750   0.4837   0.02050   0.01232  -0.1157   0.8201   0.2937
  -0.500   0.5116   0.02000   0.01190  -0.1149   0.8073   0.3251
  -0.250   0.5406   0.01936   0.01134  -0.1142   0.7952   0.3537
   0.000   0.5727   0.01852   0.01063  -0.1138   0.7845   0.3961
   0.250   0.6053   0.01650   0.00999  -0.1130   0.7722   1.0000
   0.500   0.6329   0.01638   0.00957  -0.1121   0.7574   1.0000
   0.750   0.6602   0.01624   0.00922  -0.1111   0.7421   1.0000
   1.000   0.6874   0.01609   0.00888  -0.1102   0.7261   1.0000
   1.250   0.7149   0.01591   0.00852  -0.1093   0.7092   1.0000
   1.500   0.7395   0.01586   0.00833  -0.1081   0.6885   1.0000
   1.750   0.7659   0.01574   0.00804  -0.1070   0.6673   1.0000
   2.000   0.7919   0.01569   0.00779  -0.1060   0.6437   1.0000
   2.250   0.8180   0.01571   0.00757  -0.1050   0.6182   1.0000
   2.500   0.8425   0.01590   0.00751  -0.1038   0.5891   1.0000
   2.750   0.8666   0.01625   0.00756  -0.1026   0.5585   1.0000
   3.000   0.8900   0.01675   0.00777  -0.1015   0.5279   1.0000
   3.250   0.9131   0.01733   0.00808  -0.1004   0.4993   1.0000
   3.500   0.9366   0.01790   0.00839  -0.0995   0.4753   1.0000
   3.750   0.9594   0.01841   0.00875  -0.0986   0.4535   1.0000
   4.000   0.9827   0.01889   0.00909  -0.0977   0.4356   1.0000
   4.250   1.0064   0.01938   0.00944  -0.0970   0.4204   1.0000
   4.500   1.0306   0.01989   0.00983  -0.0964   0.4077   1.0000
   4.750   1.0554   0.02042   0.01025  -0.0959   0.3973   1.0000
   5.000   1.0802   0.02095   0.01070  -0.0954   0.3874   1.0000
   5.250   1.1043   0.02151   0.01122  -0.0949   0.3777   1.0000
   5.500   1.1298   0.02208   0.01164  -0.0946   0.3693   1.0000
   5.750   1.1532   0.02266   0.01229  -0.0939   0.3607   1.0000
   6.000   1.1794   0.02331   0.01280  -0.0938   0.3538   1.0000
   6.250   1.2026   0.02394   0.01354  -0.0932   0.3464   1.0000
   6.500   1.2291   0.02462   0.01411  -0.0931   0.3400   1.0000
   6.750   1.2513   0.02532   0.01495  -0.0924   0.3330   1.0000
   7.000   1.2765   0.02600   0.01558  -0.0921   0.3262   1.0000
   7.250   1.2993   0.02678   0.01646  -0.0915   0.3195   1.0000
   7.500   1.3224   0.02749   0.01721  -0.0909   0.3123   1.0000
   7.750   1.3459   0.02834   0.01807  -0.0905   0.3055   1.0000
   8.000   1.3666   0.02911   0.01896  -0.0896   0.2980   1.0000
   8.250   1.3913   0.03003   0.01985  -0.0893   0.2913   1.0000
   8.500   1.4091   0.03092   0.02095  -0.0880   0.2837   1.0000
   8.750   1.4348   0.03186   0.02178  -0.0880   0.2767   1.0000
   9.000   1.4482   0.03264   0.02281  -0.0860   0.2680   1.0000
   9.250   1.4686   0.03333   0.02347  -0.0851   0.2597   1.0000
   9.500   1.4829   0.03391   0.02418  -0.0834   0.2513   1.0000
   9.750   1.5000   0.03463   0.02494  -0.0821   0.2438   1.0000
  10.000   1.5130   0.03519   0.02562  -0.0802   0.2363   1.0000
  10.250   1.5318   0.03590   0.02633  -0.0792   0.2303   1.0000
  10.500   1.5407   0.03686   0.02757  -0.0768   0.2244   1.0000
  10.750   1.5587   0.03728   0.02795  -0.0756   0.2189   1.0000
  11.000   1.5632   0.03808   0.02896  -0.0727   0.2128   1.0000
  11.250   1.5674   0.03858   0.02959  -0.0697   0.2068   1.0000
  11.500   1.5810   0.03901   0.02998  -0.0680   0.2021   1.0000
  11.750   1.5781   0.04028   0.03159  -0.0645   0.1978   1.0000
  12.000   1.5807   0.04122   0.03270  -0.0618   0.1931   1.0000
  12.250   1.5865   0.04172   0.03317  -0.0596   0.1882   1.0000
  12.500   1.5783   0.04342   0.03521  -0.0566   0.1828   1.0000
  12.750   1.5770   0.04472   0.03666  -0.0545   0.1774   1.0000
  13.000   1.5767   0.04624   0.03831  -0.0528   0.1728   1.0000
  13.250   1.5747   0.04831   0.04067  -0.0513   0.1678   1.0000
  13.500   1.5712   0.05030   0.04278  -0.0500   0.1620   1.0000
  13.750   1.5677   0.05285   0.04558  -0.0491   0.1565   1.0000
  14.000   1.5624   0.05561   0.04849  -0.0485   0.1504   1.0000
  14.250   1.5543   0.05898   0.05202  -0.0482   0.1434   1.0000
  14.500   1.5431   0.06290   0.05602  -0.0484   0.1369   1.0000
  14.750   1.5298   0.06756   0.06083  -0.0490   0.1301   1.0000
  15.000   1.5136   0.07284   0.06620  -0.0501   0.1247   1.0000
  15.250   1.4968   0.07864   0.07217  -0.0516   0.1200   1.0000
  15.500   1.4777   0.08505   0.07878  -0.0536   0.1160   1.0000
  15.750   1.4557   0.09209   0.08596  -0.0560   0.1129   1.0000
  16.000   1.4328   0.09949   0.09345  -0.0588   0.1102   1.0000
  16.250   1.4100   0.10723   0.10144  -0.0618   0.1081   1.0000
<< Back to GOE 363 AIRFOIL (goe363-il)

Polar data table (+)

Polar graphs


<< Back to GOE 363 AIRFOIL (goe363-il)