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

GOE 612 AIRFOIL (goe612-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 612 AIRFOIL (goe612-il)
Reynolds number: 200,000
Max Cl/Cd: 70.08 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe612-il-200000-n5.txt
Download as CSV file: xf-goe612-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 612 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
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.4703   0.04222   0.03733  -0.1276   0.9567   0.0391
 -11.000  -0.4792   0.03849   0.03330  -0.1300   0.9444   0.0393
 -10.750  -0.4773   0.03595   0.03052  -0.1304   0.9346   0.0396
 -10.500  -0.4670   0.03378   0.02809  -0.1308   0.9271   0.0400
 -10.250  -0.4602   0.03216   0.02627  -0.1294   0.9173   0.0404
 -10.000  -0.4453   0.03053   0.02439  -0.1290   0.9105   0.0410
  -9.750  -0.4351   0.02915   0.02278  -0.1274   0.9009   0.0415
  -9.500  -0.4180   0.02772   0.02108  -0.1266   0.8939   0.0421
  -9.250  -0.4034   0.02655   0.01967  -0.1251   0.8847   0.0426
  -9.000  -0.3836   0.02541   0.01829  -0.1243   0.8776   0.0430
  -8.750  -0.3638   0.02440   0.01724  -0.1234   0.8699   0.0434
  -8.500  -0.3425   0.02354   0.01633  -0.1226   0.8622   0.0439
  -8.250  -0.3196   0.02274   0.01543  -0.1220   0.8554   0.0445
  -8.000  -0.2984   0.02202   0.01463  -0.1210   0.8465   0.0451
  -7.750  -0.2738   0.02126   0.01376  -0.1205   0.8400   0.0458
  -7.500  -0.2523   0.02062   0.01302  -0.1195   0.8308   0.0466
  -7.250  -0.2274   0.01997   0.01222  -0.1190   0.8243   0.0476
  -7.000  -0.2043   0.01943   0.01155  -0.1182   0.8160   0.0486
  -6.750  -0.1802   0.01878   0.01085  -0.1176   0.8088   0.0496
  -6.500  -0.1564   0.01824   0.01028  -0.1169   0.8014   0.0506
  -6.250  -0.1320   0.01776   0.00974  -0.1163   0.7938   0.0517
  -6.000  -0.1066   0.01729   0.00918  -0.1158   0.7876   0.0530
  -5.750  -0.0825   0.01689   0.00870  -0.1150   0.7803   0.0544
  -5.500  -0.0571   0.01647   0.00820  -0.1145   0.7743   0.0558
  -5.250  -0.0323   0.01607   0.00778  -0.1140   0.7683   0.0579
  -5.000  -0.0074   0.01576   0.00742  -0.1133   0.7616   0.0606
  -4.750   0.0187   0.01544   0.00702  -0.1129   0.7561   0.0636
  -4.500   0.0441   0.01514   0.00671  -0.1124   0.7507   0.0673
  -4.250   0.0694   0.01488   0.00643  -0.1118   0.7450   0.0721
  -4.000   0.0957   0.01464   0.00618  -0.1114   0.7399   0.0790
  -3.750   0.1226   0.01444   0.00595  -0.1111   0.7352   0.0894
  -3.500   0.1479   0.01428   0.00581  -0.1106   0.7296   0.1018
  -3.250   0.1740   0.01410   0.00564  -0.1102   0.7246   0.1147
  -3.000   0.2012   0.01392   0.00543  -0.1099   0.7195   0.1277
  -2.750   0.2263   0.01377   0.00532  -0.1093   0.7129   0.1415
  -2.500   0.2519   0.01362   0.00519  -0.1088   0.7062   0.1578
  -2.000   0.3035   0.01334   0.00502  -0.1078   0.6941   0.2038
  -1.750   0.3291   0.01322   0.00497  -0.1072   0.6877   0.2336
  -1.250   0.3803   0.01306   0.00495  -0.1061   0.6752   0.2964
  -1.000   0.4062   0.01302   0.00491  -0.1055   0.6683   0.3201
  -0.750   0.4323   0.01299   0.00487  -0.1050   0.6618   0.3383
  -0.500   0.4571   0.01296   0.00488  -0.1043   0.6546   0.3550
  -0.250   0.4834   0.01295   0.00486  -0.1038   0.6489   0.3723
   0.000   0.5092   0.01294   0.00487  -0.1032   0.6434   0.3904
   0.250   0.5341   0.01293   0.00491  -0.1025   0.6373   0.4079
   0.500   0.5599   0.01292   0.00490  -0.1020   0.6313   0.4258
   0.750   0.5850   0.01290   0.00493  -0.1013   0.6251   0.4456
   1.000   0.6090   0.01287   0.00498  -0.1004   0.6178   0.4712
   1.250   0.6337   0.01281   0.00498  -0.0996   0.6114   0.5034
   1.500   0.6563   0.01272   0.00506  -0.0985   0.6041   0.5406
   1.750   0.6784   0.01256   0.00509  -0.0972   0.5964   0.5969
   2.250   0.7994   0.01210   0.00534  -0.1103   0.5750   1.0000
   2.500   0.8206   0.01223   0.00543  -0.1089   0.5647   1.0000
   2.750   0.8416   0.01236   0.00548  -0.1075   0.5536   1.0000
   3.000   0.8618   0.01251   0.00554  -0.1058   0.5408   1.0000
   3.250   0.8813   0.01267   0.00563  -0.1041   0.5258   1.0000
   3.500   0.8999   0.01286   0.00573  -0.1022   0.5092   1.0000
   3.750   0.9173   0.01309   0.00583  -0.1001   0.4904   1.0000
   4.000   0.9333   0.01335   0.00598  -0.0978   0.4692   1.0000
   4.250   0.9475   0.01368   0.00616  -0.0951   0.4467   1.0000
   4.500   0.9601   0.01406   0.00638  -0.0923   0.4256   1.0000
   4.750   0.9719   0.01447   0.00665  -0.0893   0.4068   1.0000
   5.000   0.9826   0.01488   0.00694  -0.0861   0.3907   1.0000
   5.250   0.9940   0.01530   0.00726  -0.0831   0.3768   1.0000
   5.500   1.0064   0.01573   0.00760  -0.0804   0.3650   1.0000
   5.750   1.0183   0.01622   0.00799  -0.0777   0.3540   1.0000
   6.000   1.0328   0.01666   0.00837  -0.0755   0.3433   1.0000
   6.250   1.0464   0.01716   0.00879  -0.0732   0.3341   1.0000
   6.500   1.0617   0.01763   0.00922  -0.0713   0.3253   1.0000
   6.750   1.0760   0.01815   0.00968  -0.0693   0.3181   1.0000
   7.000   1.0932   0.01858   0.01011  -0.0677   0.3113   1.0000
   7.250   1.1088   0.01909   0.01058  -0.0660   0.3053   1.0000
   7.500   1.1246   0.01961   0.01108  -0.0643   0.3001   1.0000
   7.750   1.1423   0.02006   0.01155  -0.0629   0.2946   1.0000
   8.000   1.1581   0.02060   0.01208  -0.0613   0.2891   1.0000
   8.250   1.1727   0.02121   0.01265  -0.0596   0.2841   1.0000
   8.500   1.1902   0.02170   0.01320  -0.0584   0.2785   1.0000
   8.750   1.2059   0.02228   0.01379  -0.0569   0.2731   1.0000
   9.000   1.2195   0.02297   0.01445  -0.0552   0.2679   1.0000
   9.250   1.2358   0.02354   0.01509  -0.0540   0.2626   1.0000
   9.500   1.2509   0.02419   0.01576  -0.0526   0.2570   1.0000
   9.750   1.2640   0.02495   0.01651  -0.0510   0.2520   1.0000
  10.000   1.2789   0.02564   0.01725  -0.0497   0.2471   1.0000
  10.250   1.2932   0.02638   0.01805  -0.0484   0.2415   1.0000
  10.500   1.3054   0.02725   0.01891  -0.0470   0.2364   1.0000
  10.750   1.3185   0.02810   0.01980  -0.0457   0.2315   1.0000
  11.000   1.3318   0.02895   0.02071  -0.0444   0.2258   1.0000
  11.250   1.3425   0.02998   0.02175  -0.0430   0.2206   1.0000
  11.500   1.3542   0.03097   0.02279  -0.0418   0.2154   1.0000
  11.750   1.3657   0.03201   0.02388  -0.0406   0.2096   1.0000
  12.000   1.3744   0.03325   0.02512  -0.0393   0.2044   1.0000
  12.250   1.3854   0.03438   0.02631  -0.0382   0.1993   1.0000
  12.500   1.3956   0.03560   0.02758  -0.0371   0.1943   1.0000
  12.750   1.4033   0.03701   0.02901  -0.0359   0.1898   1.0000
  13.000   1.4130   0.03832   0.03037  -0.0349   0.1856   1.0000
  13.250   1.4224   0.03968   0.03180  -0.0340   0.1811   1.0000
  13.500   1.4296   0.04124   0.03339  -0.0330   0.1769   1.0000
  13.750   1.4358   0.04291   0.03507  -0.0320   0.1730   1.0000
  14.000   1.4449   0.04439   0.03664  -0.0313   0.1688   1.0000
  14.250   1.4509   0.04616   0.03846  -0.0305   0.1642   1.0000
  14.500   1.4544   0.04816   0.04047  -0.0297   0.1604   1.0000
  14.750   1.4619   0.04987   0.04227  -0.0290   0.1564   1.0000
  15.000   1.4674   0.05180   0.04428  -0.0285   0.1524   1.0000
  15.250   1.4712   0.05392   0.04643  -0.0279   0.1491   1.0000
  15.500   1.4738   0.05618   0.04869  -0.0274   0.1461   1.0000
  15.750   1.4808   0.05806   0.05071  -0.0270   0.1435   1.0000
  16.000   1.4858   0.06016   0.05292  -0.0267   0.1406   1.0000
  16.250   1.4882   0.06260   0.05542  -0.0264   0.1374   1.0000
  16.500   1.4898   0.06513   0.05800  -0.0263   0.1349   1.0000
  16.750   1.4912   0.06772   0.06063  -0.0261   0.1324   1.0000
  17.000   1.4945   0.07021   0.06327  -0.0261   0.1296   1.0000
  17.250   1.4962   0.07289   0.06607  -0.0262   0.1270   1.0000
  17.500   1.4965   0.07578   0.06905  -0.0264   0.1246   1.0000
  17.750   1.4947   0.07897   0.07229  -0.0267   0.1221   1.0000
  18.000   1.4932   0.08214   0.07551  -0.0271   0.1199   1.0000
  18.250   1.4919   0.08545   0.07899  -0.0277   0.1170   1.0000
  18.500   1.4891   0.08898   0.08266  -0.0284   0.1143   1.0000
  18.750   1.4853   0.09267   0.08644  -0.0292   0.1119   1.0000
  19.000   1.4801   0.09661   0.09045  -0.0303   0.1095   1.0000
  19.250   1.4744   0.10069   0.09463  -0.0314   0.1069   1.0000
<< Back to GOE 612 AIRFOIL (goe612-il)

Polar data table (+)

Polar graphs


<< Back to GOE 612 AIRFOIL (goe612-il)