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

GOE 562 AIRFOIL (goe562-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 562 AIRFOIL (goe562-il)
Reynolds number: 500,000
Max Cl/Cd: 101.4 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe562-il-500000.txt
Download as CSV file: xf-goe562-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 562 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500   0.0263   0.08689   0.08458  -0.0919   0.8884   0.0236
 -10.250   0.0620   0.07930   0.07680  -0.1055   0.8670   0.0258
 -10.000  -0.1473   0.10123   0.09911  -0.0616   0.9211   0.0222
  -9.750  -0.1010   0.09516   0.09294  -0.0732   0.9069   0.0227
  -9.500  -0.0362   0.08731   0.08492  -0.0918   0.8899   0.0246
  -9.250   0.0094   0.07734   0.07468  -0.1166   0.8552   0.0262
  -9.000   0.0435   0.07206   0.06909  -0.1246   0.8048   0.0266
  -8.750   0.0601   0.07001   0.06680  -0.1259   0.7621   0.0269
  -8.500   0.0688   0.06789   0.06452  -0.1268   0.7314   0.0274
  -8.250   0.0743   0.06556   0.06207  -0.1279   0.7088   0.0280
  -8.000   0.0750   0.06284   0.05927  -0.1296   0.6917   0.0289
  -7.750   0.0702   0.06003   0.05640  -0.1305   0.6781   0.0294
  -7.500   0.0622   0.05694   0.05324  -0.1309   0.6676   0.0304
  -7.250   0.0489   0.05090   0.04698  -0.1335   0.6601   0.0316
  -7.000   0.0379   0.04600   0.04191  -0.1315   0.6530   0.0320
  -6.750   0.0456   0.04419   0.04005  -0.1300   0.6441   0.0323
  -6.500   0.0554   0.04282   0.03863  -0.1285   0.6360   0.0327
  -6.250   0.0654   0.04132   0.03705  -0.1269   0.6286   0.0333
  -6.000   0.0751   0.03952   0.03513  -0.1252   0.6222   0.0340
  -5.750   0.0843   0.03745   0.03294  -0.1232   0.6159   0.0354
  -5.500   0.0911   0.03542   0.03031  -0.1188   0.6109   0.0384
  -5.250   0.0893   0.03088   0.02562  -0.1155   0.6063   0.0393
  -5.000   0.1047   0.02960   0.02434  -0.1140   0.6003   0.0400
  -4.750   0.1206   0.02855   0.02319  -0.1124   0.5947   0.0409
  -4.500   0.1358   0.02735   0.02186  -0.1103   0.5893   0.0423
  -4.250   0.1559   0.02750   0.02164  -0.1075   0.5835   0.0466
  -4.000   0.1584   0.02434   0.01808  -0.1030   0.5792   0.0478
  -3.750   0.1767   0.02276   0.01649  -0.1018   0.5744   0.0489
  -3.500   0.1953   0.02184   0.01554  -0.1002   0.5691   0.0503
  -3.250   0.2147   0.02111   0.01469  -0.0986   0.5641   0.0528
  -3.000   0.2406   0.02224   0.01554  -0.0972   0.5592   0.0571
  -2.750   0.2503   0.01919   0.01228  -0.0941   0.5547   0.0597
  -2.500   0.2710   0.01832   0.01138  -0.0928   0.5493   0.0615
  -2.250   0.2928   0.01779   0.01071  -0.0915   0.5441   0.0643
  -2.000   0.3184   0.01561   0.00800  -0.0892   0.5399   0.0454
  -1.750   0.3420   0.01478   0.00706  -0.0881   0.5344   0.0448
  -1.500   0.3667   0.01425   0.00640  -0.0873   0.5295   0.0448
  -1.250   0.3910   0.01375   0.00584  -0.0864   0.5246   0.0449
  -1.000   0.4144   0.01349   0.00553  -0.0854   0.5192   0.0458
  -0.750   0.4385   0.01326   0.00521  -0.0845   0.5145   0.0465
  -0.500   0.4623   0.01297   0.00490  -0.0836   0.5101   0.0467
  -0.250   0.4849   0.01259   0.00451  -0.0824   0.5057   0.0473
   0.000   0.5055   0.01217   0.00406  -0.0808   0.5013   0.0486
   0.250   0.5273   0.01201   0.00386  -0.0795   0.4967   0.0500
   0.500   0.5488   0.01185   0.00372  -0.0782   0.4923   0.0520
   0.750   0.5705   0.01175   0.00361  -0.0768   0.4876   0.0545
   1.000   0.5924   0.01171   0.00351  -0.0756   0.4831   0.0561
   1.250   0.6142   0.01161   0.00340  -0.0742   0.4787   0.0590
   1.500   0.6355   0.01150   0.00330  -0.0728   0.4745   0.0652
   1.750   0.6568   0.01144   0.00324  -0.0714   0.4705   0.0803
   2.000   0.9338   0.01048   0.00432  -0.1282   0.4581   1.0000
   2.250   0.9553   0.01059   0.00437  -0.1270   0.4539   1.0000
   2.500   0.9770   0.01072   0.00445  -0.1258   0.4500   1.0000
   2.750   0.9992   0.01078   0.00453  -0.1247   0.4460   1.0000
   3.000   1.0207   0.01088   0.00460  -0.1235   0.4417   1.0000
   3.250   1.0419   0.01102   0.00469  -0.1222   0.4378   1.0000
   3.500   1.0633   0.01115   0.00479  -0.1210   0.4339   1.0000
   3.750   1.0850   0.01123   0.00490  -0.1199   0.4298   1.0000
   4.000   1.1059   0.01135   0.00500  -0.1185   0.4255   1.0000
   4.250   1.1260   0.01152   0.00512  -0.1171   0.4212   1.0000
   4.500   1.1471   0.01163   0.00525  -0.1158   0.4171   1.0000
   4.750   1.1677   0.01173   0.00538  -0.1145   0.4124   1.0000
   5.000   1.1871   0.01189   0.00552  -0.1129   0.4079   1.0000
   5.250   1.2065   0.01205   0.00567  -0.1113   0.4034   1.0000
   5.500   1.2265   0.01216   0.00582  -0.1098   0.3983   1.0000
   5.750   1.2447   0.01232   0.00597  -0.1080   0.3931   1.0000
   6.000   1.2617   0.01249   0.00614  -0.1060   0.3880   1.0000
   6.250   1.2786   0.01261   0.00630  -0.1038   0.3821   1.0000
   6.500   1.2920   0.01281   0.00647  -0.1011   0.3756   1.0000
   6.750   1.3078   0.01294   0.00663  -0.0988   0.3670   1.0000
   7.000   1.3208   0.01317   0.00683  -0.0960   0.3582   1.0000
   7.250   1.3333   0.01342   0.00705  -0.0931   0.3463   1.0000
   7.500   1.3468   0.01368   0.00730  -0.0905   0.3345   1.0000
   7.750   1.3591   0.01400   0.00759  -0.0877   0.3224   1.0000
   8.000   1.3703   0.01438   0.00793  -0.0848   0.3108   1.0000
   8.250   1.3804   0.01481   0.00833  -0.0817   0.2995   1.0000
   8.500   1.3902   0.01529   0.00876  -0.0787   0.2878   1.0000
   8.750   1.4016   0.01573   0.00921  -0.0760   0.2791   1.0000
   9.000   1.4107   0.01628   0.00973  -0.0730   0.2714   1.0000
   9.250   1.4206   0.01681   0.01026  -0.0702   0.2623   1.0000
   9.500   1.4292   0.01743   0.01086  -0.0673   0.2544   1.0000
   9.750   1.4393   0.01801   0.01146  -0.0647   0.2483   1.0000
  10.000   1.4467   0.01873   0.01218  -0.0619   0.2414   1.0000
  10.250   1.4553   0.01944   0.01291  -0.0594   0.2339   1.0000
  10.500   1.4617   0.02031   0.01379  -0.0566   0.2272   1.0000
  10.750   1.4650   0.02140   0.01486  -0.0538   0.2166   1.0000
  11.000   1.4698   0.02250   0.01597  -0.0513   0.2053   1.0000
  11.250   1.4685   0.02406   0.01749  -0.0484   0.1893   1.0000
  11.500   1.4500   0.02697   0.02019  -0.0444   0.1450   1.0000
  11.750   1.4200   0.03117   0.02422  -0.0404   0.1071   1.0000
  12.000   1.3911   0.03582   0.02880  -0.0375   0.0742   1.0000
  12.250   1.3741   0.03970   0.03265  -0.0357   0.0572   1.0000
  12.500   1.3611   0.04337   0.03635  -0.0344   0.0440   1.0000
  12.750   1.3458   0.04739   0.04036  -0.0332   0.0294   1.0000
  13.000   1.3356   0.05100   0.04398  -0.0324   0.0240   1.0000
  13.250   1.3271   0.05454   0.04759  -0.0317   0.0216   1.0000
  13.500   1.3201   0.05801   0.05114  -0.0312   0.0201   1.0000
  13.750   1.3146   0.06138   0.05461  -0.0308   0.0195   1.0000
  14.000   1.3086   0.06489   0.05822  -0.0305   0.0189   1.0000
  14.250   1.3032   0.06838   0.06181  -0.0303   0.0186   1.0000
  14.500   1.2932   0.07252   0.06605  -0.0303   0.0178   1.0000
  14.750   1.2836   0.07671   0.07035  -0.0303   0.0175   1.0000
  15.000   1.2734   0.08108   0.07481  -0.0305   0.0170   1.0000
  15.250   1.2624   0.08559   0.07942  -0.0308   0.0168   1.0000
<< Back to GOE 562 AIRFOIL (goe562-il)

Polar data table (+)

Polar graphs


<< Back to GOE 562 AIRFOIL (goe562-il)