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

GOE 361 AIRFOIL (goe361-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 361 AIRFOIL (goe361-il)
Reynolds number: 200,000
Max Cl/Cd: 76.64 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe361-il-200000.txt
Download as CSV file: xf-goe361-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 361 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.000  -0.2852   0.08925   0.08611  -0.0214   1.0000   0.0281
  -6.750  -0.2859   0.08715   0.08406  -0.0202   1.0000   0.0285
  -6.500  -0.2888   0.08532   0.08231  -0.0189   1.0000   0.0290
  -6.250  -0.2926   0.08360   0.08065  -0.0176   1.0000   0.0297
  -6.000  -0.2962   0.08193   0.07903  -0.0166   1.0000   0.0301
  -5.750  -0.2899   0.07955   0.07668  -0.0182   0.9988   0.0312
  -5.500  -0.2187   0.07460   0.07157  -0.0409   0.9904   0.0345
  -5.250  -0.1716   0.06815   0.06502  -0.0532   0.9845   0.0351
  -5.000  -0.1559   0.06364   0.06056  -0.0527   0.9785   0.0361
  -4.750  -0.1225   0.05975   0.05663  -0.0575   0.9728   0.0379
  -4.500  -0.0807   0.05572   0.05253  -0.0653   0.9655   0.0412
  -4.250  -0.0087   0.04964   0.04612  -0.0822   0.9592   0.0458
  -4.000   0.0179   0.04627   0.04278  -0.0841   0.9526   0.0477
  -3.750   0.0589   0.04299   0.03942  -0.0895   0.9462   0.0517
  -3.500   0.1188   0.03837   0.03444  -0.0996   0.9406   0.0587
  -3.250   0.1493   0.03589   0.03196  -0.1017   0.9310   0.0623
  -3.000   0.2012   0.03302   0.02857  -0.1079   0.9206   0.0715
  -2.750   0.2299   0.03023   0.02588  -0.1095   0.9091   0.0746
  -2.500   0.2709   0.02846   0.02362  -0.1123   0.8945   0.0858
  -2.250   0.2947   0.02624   0.02147  -0.1125   0.8759   0.0900
  -2.000   0.3255   0.02462   0.01953  -0.1131   0.8557   0.1019
  -1.500   0.3775   0.02188   0.01644  -0.1126   0.8117   0.1204
  -1.250   0.4040   0.02071   0.01504  -0.1124   0.7916   0.1342
  -1.000   0.4303   0.01971   0.01384  -0.1121   0.7720   0.1510
  -0.500   0.4805   0.01791   0.01173  -0.1113   0.7351   0.2058
  -0.250   0.5225   0.01550   0.00814  -0.1099   0.7200   0.0797
   0.000   0.5504   0.01481   0.00705  -0.1088   0.7022   0.0694
   0.250   0.5767   0.01400   0.00605  -0.1082   0.6845   0.0669
   0.500   0.6029   0.01347   0.00536  -0.1074   0.6670   0.0653
   0.750   0.6288   0.01309   0.00484  -0.1066   0.6499   0.0647
   1.000   0.6545   0.01281   0.00446  -0.1059   0.6331   0.0653
   1.250   0.6802   0.01262   0.00421  -0.1052   0.6156   0.0671
   1.500   0.7058   0.01253   0.00407  -0.1045   0.5980   0.0710
   1.750   0.7313   0.01237   0.00386  -0.1039   0.5804   0.0744
   2.000   0.7569   0.01236   0.00378  -0.1032   0.5627   0.0800
   2.500   0.8145   0.01085   0.00380  -0.1036   0.5263   1.0000
   2.750   0.8398   0.01108   0.00385  -0.1030   0.5078   1.0000
   3.000   0.8649   0.01134   0.00396  -0.1023   0.4896   1.0000
   3.250   0.8898   0.01163   0.00409  -0.1016   0.4719   1.0000
   3.500   0.9147   0.01194   0.00427  -0.1010   0.4543   1.0000
   3.750   0.9396   0.01226   0.00448  -0.1004   0.4374   1.0000
   4.000   0.9644   0.01261   0.00475  -0.0997   0.4216   1.0000
   4.250   0.9891   0.01299   0.00504  -0.0992   0.4071   1.0000
   4.500   1.0139   0.01338   0.00535  -0.0986   0.3941   1.0000
   5.000   1.0639   0.01415   0.00606  -0.0975   0.3720   1.0000
   5.250   1.0890   0.01457   0.00646  -0.0970   0.3632   1.0000
   5.500   1.1134   0.01493   0.00676  -0.0964   0.3525   1.0000
   5.750   1.1371   0.01520   0.00700  -0.0957   0.3392   1.0000
   6.000   1.1606   0.01542   0.00728  -0.0950   0.3257   1.0000
   6.250   1.1844   0.01566   0.00758  -0.0943   0.3133   1.0000
   6.500   1.2079   0.01593   0.00790  -0.0936   0.3005   1.0000
   6.750   1.2303   0.01619   0.00817  -0.0927   0.2833   1.0000
   7.000   1.2520   0.01647   0.00846  -0.0917   0.2558   1.0000
   7.250   1.2716   0.01707   0.00886  -0.0904   0.2048   1.0000
   7.750   1.2892   0.02128   0.01202  -0.0854   0.0399   1.0000
   8.000   1.3049   0.02248   0.01341  -0.0836   0.0365   1.0000
   8.250   1.3204   0.02356   0.01471  -0.0817   0.0347   1.0000
   8.500   1.3325   0.02481   0.01612  -0.0795   0.0325   1.0000
   8.750   1.3406   0.02626   0.01768  -0.0769   0.0308   1.0000
   9.000   1.3450   0.02786   0.01939  -0.0738   0.0298   1.0000
   9.250   1.3447   0.02956   0.02116  -0.0701   0.0291   1.0000
   9.500   1.3433   0.03149   0.02316  -0.0665   0.0285   1.0000
   9.750   1.3450   0.03354   0.02526  -0.0636   0.0281   1.0000
  10.000   1.3522   0.03588   0.02761  -0.0613   0.0277   1.0000
  10.250   1.3696   0.03829   0.03008  -0.0600   0.0271   1.0000
  10.500   1.3792   0.03969   0.03163  -0.0582   0.0265   1.0000
  10.750   1.3911   0.04140   0.03349  -0.0566   0.0258   1.0000
  11.000   1.4118   0.04371   0.03594  -0.0557   0.0257   1.0000
  11.250   1.4368   0.04665   0.03905  -0.0552   0.0259   1.0000
  11.500   1.4992   0.05229   0.04485  -0.0595   0.0279   1.0000
  11.750   1.4944   0.05262   0.04544  -0.0551   0.0289   1.0000
<< Back to GOE 361 AIRFOIL (goe361-il)

Polar data table (+)

Polar graphs


<< Back to GOE 361 AIRFOIL (goe361-il)