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

GOE 367 AIRFOIL (goe367-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 367 AIRFOIL (goe367-il)
Reynolds number: 50,000
Max Cl/Cd: 20.82 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe367-il-50000-n5.txt
Download as CSV file: xf-goe367-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 367 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.3047   0.13583   0.12870  -0.0399   1.0000   0.1385
 -10.750  -0.3135   0.13444   0.12739  -0.0382   1.0000   0.1403
 -10.500  -0.3215   0.13268   0.12569  -0.0376   0.9992   0.1423
 -10.000  -0.3186   0.11939   0.11228  -0.0490   0.9886   0.1006
  -9.500  -0.3274   0.10730   0.10016  -0.0602   0.9759   0.0913
  -9.250  -0.3198   0.10291   0.09574  -0.0640   0.9703   0.0910
  -9.000  -0.3196   0.09841   0.09124  -0.0676   0.9629   0.0909
  -8.750  -0.3207   0.09340   0.08621  -0.0723   0.9559   0.0907
  -8.500  -0.3295   0.08917   0.08196  -0.0748   0.9465   0.0904
  -8.250  -0.3317   0.08415   0.07686  -0.0785   0.9391   0.0901
  -8.000  -0.3409   0.08019   0.07283  -0.0790   0.9287   0.0897
  -7.750  -0.3396   0.07516   0.06763  -0.0820   0.9222   0.0893
  -7.500  -0.3496   0.07164   0.06397  -0.0809   0.9116   0.0890
  -7.250  -0.3448   0.06701   0.05905  -0.0827   0.9054   0.0890
  -7.000  -0.3512   0.06378   0.05558  -0.0807   0.8954   0.0895
  -6.750  -0.3448   0.05938   0.05069  -0.0814   0.8890   0.0907
  -6.500  -0.3439   0.05629   0.04719  -0.0795   0.8804   0.0917
  -6.250  -0.3235   0.05465   0.04550  -0.0796   0.8741   0.0931
  -6.000  -0.2976   0.05219   0.04274  -0.0809   0.8698   0.0945
  -5.750  -0.2926   0.05058   0.04091  -0.0781   0.8602   0.0954
  -5.500  -0.2690   0.04845   0.03844  -0.0785   0.8548   0.0974
  -5.250  -0.2444   0.04634   0.03588  -0.0789   0.8497   0.1005
  -5.000  -0.2338   0.04476   0.03381  -0.0766   0.8406   0.1029
  -4.750  -0.2025   0.04334   0.03229  -0.0778   0.8360   0.1053
  -4.500  -0.1817   0.04236   0.03117  -0.0770   0.8289   0.1077
  -4.250  -0.1592   0.04141   0.02999  -0.0764   0.8216   0.1116
  -4.000  -0.1243   0.04000   0.02815  -0.0776   0.8174   0.1168
  -3.750  -0.1071   0.03945   0.02760  -0.0762   0.8088   0.1199
  -3.500  -0.0783   0.03864   0.02666  -0.0764   0.8026   0.1252
  -3.250  -0.0403   0.03760   0.02538  -0.0780   0.7987   0.1329
  -3.000  -0.0260   0.03735   0.02515  -0.0760   0.7889   0.1378
  -2.750   0.0081   0.03660   0.02418  -0.0768   0.7834   0.1476
  -2.500   0.0485   0.03570   0.02328  -0.0787   0.7800   0.1598
  -2.250   0.0622   0.03566   0.02325  -0.0766   0.7692   0.1685
  -2.000   0.0993   0.03495   0.02249  -0.0779   0.7643   0.1849
  -1.750   0.1256   0.03459   0.02212  -0.0776   0.7574   0.2014
  -1.500   0.1464   0.03437   0.02196  -0.0766   0.7487   0.2197
  -1.250   0.1815   0.03356   0.02131  -0.0776   0.7444   0.2500
  -1.000   0.1927   0.03366   0.02154  -0.0751   0.7334   0.2746
  -0.750   0.2219   0.03275   0.02103  -0.0752   0.7279   0.3320
  -0.250   0.4166   0.03082   0.02068  -0.0975   0.7216   1.0000
   0.000   0.4283   0.03110   0.02079  -0.0948   0.7113   1.0000
   0.250   0.4635   0.03067   0.02011  -0.0956   0.7067   1.0000
   0.500   0.4612   0.03142   0.02078  -0.0908   0.6936   1.0000
   0.750   0.4962   0.03098   0.02013  -0.0915   0.6889   1.0000
   1.000   0.4929   0.03180   0.02088  -0.0867   0.6760   1.0000
   1.250   0.5274   0.03140   0.02030  -0.0872   0.6713   1.0000
   1.500   0.5233   0.03229   0.02113  -0.0823   0.6587   1.0000
   1.750   0.5573   0.03191   0.02059  -0.0828   0.6540   1.0000
   2.000   0.5529   0.03285   0.02149  -0.0779   0.6416   1.0000
   2.250   0.5861   0.03251   0.02101  -0.0783   0.6368   1.0000
   2.500   0.5810   0.03355   0.02200  -0.0734   0.6248   1.0000
   2.750   0.6134   0.03326   0.02159  -0.0737   0.6200   1.0000
   3.000   0.6080   0.03434   0.02263  -0.0688   0.6085   1.0000
   3.250   0.6392   0.03414   0.02234  -0.0690   0.6034   1.0000
   3.500   0.6371   0.03523   0.02339  -0.0648   0.5929   1.0000
   3.750   0.6640   0.03521   0.02330  -0.0644   0.5872   1.0000
   4.000   0.7019   0.03480   0.02279  -0.0654   0.5835   1.0000
   4.250   0.6899   0.03649   0.02449  -0.0603   0.5713   1.0000
   4.500   0.7261   0.03612   0.02404  -0.0610   0.5677   1.0000
   4.750   0.7157   0.03810   0.02604  -0.0566   0.5562   1.0000
   5.000   0.7460   0.03801   0.02590  -0.0567   0.5520   1.0000
   5.250   0.7832   0.03761   0.02545  -0.0575   0.5488   1.0000
   5.500   0.7653   0.04027   0.02816  -0.0528   0.5364   1.0000
   5.750   0.7997   0.03996   0.02781  -0.0532   0.5332   1.0000
   6.250   0.8085   0.04332   0.03121  -0.0491   0.5174   1.0000
   6.500   0.8420   0.04310   0.03098  -0.0494   0.5149   1.0000
   7.000   0.8397   0.04780   0.03576  -0.0453   0.4987   1.0000
   7.250   0.8602   0.04855   0.03653  -0.0447   0.4946   1.0000
   7.750   0.8582   0.05402   0.04211  -0.0417   0.4800   1.0000
   8.750   0.8196   0.07057   0.05890  -0.0379   0.4487   1.0000
   9.000   0.8381   0.07178   0.06017  -0.0376   0.4458   1.0000
   9.250   0.8636   0.07218   0.06064  -0.0373   0.4438   1.0000
   9.500   0.8225   0.07993   0.06846  -0.0366   0.4322   1.0000
   9.750   0.8417   0.08100   0.06960  -0.0363   0.4288   1.0000
  10.000   0.8724   0.08065   0.06931  -0.0358   0.4264   1.0000
  10.250   0.8432   0.08672   0.07544  -0.0352   0.4143   1.0000
  10.500   0.8730   0.08623   0.07504  -0.0346   0.4105   1.0000
  10.750   0.8567   0.09083   0.07970  -0.0341   0.3997   1.0000
  11.000   0.8786   0.09126   0.08022  -0.0335   0.3950   1.0000
  11.250   0.9098   0.09063   0.07967  -0.0329   0.3922   1.0000
  11.750   0.9085   0.09669   0.08592  -0.0322   0.3765   1.0000
  12.000   0.8923   0.10155   0.09087  -0.0323   0.3656   1.0000
  12.250   0.9190   0.10097   0.09039  -0.0314   0.3602   1.0000
  12.500   0.9100   0.10492   0.09443  -0.0314   0.3496   1.0000
  12.750   0.9326   0.10478   0.09439  -0.0306   0.3435   1.0000
  13.250   0.9459   0.10867   0.09850  -0.0299   0.3266   1.0000
  13.750   0.9638   0.11166   0.10170  -0.0292   0.3091   1.0000
  14.000   0.9560   0.11563   0.10577  -0.0297   0.2971   1.0000
  14.250   0.9826   0.11442   0.10470  -0.0285   0.2915   1.0000
  14.500   0.9741   0.11851   0.10888  -0.0292   0.2784   1.0000
<< Back to GOE 367 AIRFOIL (goe367-il)

Polar data table (+)

Polar graphs


<< Back to GOE 367 AIRFOIL (goe367-il)