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

GOE 481A AIRFOIL (goe481a-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 481A AIRFOIL (goe481a-il)
Reynolds number: 500,000
Max Cl/Cd: 74.08 at α=6.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe481a-il-500000-n5.txt
Download as CSV file: xf-goe481a-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 481A 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.250  -0.6223   0.03953   0.03570  -0.1404   0.9500   0.0432
 -12.000  -0.6449   0.03465   0.03048  -0.1406   0.9461   0.0434
 -11.750  -0.6600   0.03276   0.02842  -0.1357   0.9387   0.0436
 -11.500  -0.6539   0.03074   0.02620  -0.1344   0.9356   0.0438
 -11.250  -0.6399   0.02899   0.02425  -0.1339   0.9338   0.0441
 -11.000  -0.6203   0.02744   0.02252  -0.1339   0.9325   0.0443
 -10.750  -0.6185   0.02653   0.02148  -0.1298   0.9259   0.0445
 -10.500  -0.5992   0.02537   0.02017  -0.1290   0.9223   0.0447
 -10.250  -0.5737   0.02419   0.01882  -0.1293   0.9200   0.0449
 -10.000  -0.5476   0.02302   0.01756  -0.1296   0.9184   0.0452
  -9.750  -0.5192   0.02213   0.01663  -0.1301   0.9173   0.0456
  -9.500  -0.5143   0.02166   0.01611  -0.1258   0.9089   0.0458
  -9.250  -0.4898   0.02096   0.01536  -0.1253   0.9062   0.0461
  -9.000  -0.4640   0.02028   0.01463  -0.1251   0.9045   0.0464
  -8.750  -0.4370   0.01963   0.01391  -0.1250   0.9032   0.0467
  -8.500  -0.4085   0.01896   0.01317  -0.1253   0.9020   0.0470
  -8.250  -0.3981   0.01863   0.01279  -0.1218   0.8957   0.0473
  -8.000  -0.3767   0.01813   0.01223  -0.1206   0.8918   0.0476
  -7.750  -0.3506   0.01757   0.01160  -0.1202   0.8894   0.0480
  -7.500  -0.3220   0.01699   0.01094  -0.1203   0.8873   0.0483
  -7.250  -0.2903   0.01640   0.01027  -0.1210   0.8855   0.0488
  -7.000  -0.2748   0.01610   0.00991  -0.1184   0.8792   0.0491
  -6.750  -0.2533   0.01574   0.00948  -0.1170   0.8740   0.0495
  -6.500  -0.2235   0.01527   0.00894  -0.1172   0.8707   0.0498
  -6.250  -0.1883   0.01477   0.00837  -0.1186   0.8680   0.0501
  -6.000  -0.1740   0.01445   0.00803  -0.1157   0.8597   0.0505
  -5.750  -0.1457   0.01403   0.00759  -0.1156   0.8538   0.0511
  -5.500  -0.1084   0.01357   0.00711  -0.1174   0.8496   0.0518
  -5.250  -0.0902   0.01337   0.00687  -0.1151   0.8393   0.0521
  -5.000  -0.0524   0.01294   0.00640  -0.1170   0.8317   0.0528
  -4.750  -0.0260   0.01267   0.00607  -0.1164   0.8184   0.0534
  -4.500   0.0069   0.01235   0.00568  -0.1172   0.8040   0.0540
  -4.250   0.0405   0.01207   0.00528  -0.1181   0.7831   0.0546
  -4.000   0.0707   0.01190   0.00494  -0.1183   0.7546   0.0552
  -3.750   0.0917   0.01189   0.00472  -0.1166   0.7207   0.0556
  -3.500   0.1064   0.01188   0.00456  -0.1136   0.6883   0.0563
  -3.250   0.1210   0.01191   0.00445  -0.1106   0.6630   0.0569
  -3.000   0.1375   0.01193   0.00437  -0.1080   0.6432   0.0576
  -2.750   0.1556   0.01193   0.00429  -0.1058   0.6270   0.0585
  -2.500   0.1744   0.01193   0.00422  -0.1037   0.6134   0.0593
  -2.000   0.2124   0.01195   0.00410  -0.0996   0.5878   0.0611
  -1.750   0.2314   0.01195   0.00403  -0.0975   0.5772   0.0620
  -1.500   0.2524   0.01190   0.00397  -0.0959   0.5678   0.0632
  -1.000   0.2941   0.01189   0.00389  -0.0926   0.5507   0.0660
  -0.750   0.3144   0.01191   0.00386  -0.0908   0.5413   0.0675
  -0.500   0.3360   0.01190   0.00383  -0.0893   0.5340   0.0696
  -0.250   0.3576   0.01189   0.00382  -0.0879   0.5264   0.0726
   0.000   0.3780   0.01191   0.00382  -0.0862   0.5190   0.0766
   0.500   0.4214   0.01189   0.00386  -0.0833   0.5055   0.0925
   0.750   0.4416   0.01197   0.00394  -0.0816   0.4984   0.1033
   1.000   0.4641   0.01202   0.00402  -0.0803   0.4914   0.1127
   1.250   0.4834   0.01215   0.00409  -0.0784   0.4802   0.1188
   1.500   0.5054   0.01224   0.00418  -0.0771   0.4728   0.1246
   1.750   0.5269   0.01234   0.00424  -0.0756   0.4644   0.1286
   2.000   0.5474   0.01245   0.00433  -0.0740   0.4576   0.1322
   2.250   0.5702   0.01252   0.00441  -0.0728   0.4511   0.1359
   2.500   0.5915   0.01263   0.00449  -0.0714   0.4441   0.1388
   2.750   0.6119   0.01276   0.00458  -0.0698   0.4376   0.1411
   3.000   0.6336   0.01282   0.00466  -0.0684   0.4308   0.1442
   3.250   0.6528   0.01297   0.00479  -0.0666   0.4218   0.1481
   3.500   0.6737   0.01309   0.00489  -0.0651   0.4135   0.1514
   3.750   0.6928   0.01328   0.00503  -0.0634   0.4029   0.1541
   4.000   0.7135   0.01338   0.00513  -0.0619   0.3960   0.1577
   4.250   0.7341   0.01350   0.00526  -0.0604   0.3885   0.1612
   4.500   0.7531   0.01368   0.00541  -0.0587   0.3816   0.1639
   4.750   0.7745   0.01380   0.00554  -0.0574   0.3757   0.1665
   5.000   0.7947   0.01397   0.00569  -0.0559   0.3691   0.1687
   5.250   0.8138   0.01417   0.00587  -0.0542   0.3619   0.1714
   5.500   0.8345   0.01433   0.00604  -0.0529   0.3544   0.1752
   5.750   0.8536   0.01455   0.00625  -0.0513   0.3472   0.1794
   6.000   0.8736   0.01475   0.00645  -0.0498   0.3417   0.1841
   6.250   0.8940   0.01493   0.00667  -0.0485   0.3355   0.1941
   6.750   1.1090   0.01497   0.00840  -0.0845   0.3109   1.0000
   7.000   1.1280   0.01528   0.00868  -0.0830   0.3045   1.0000
   7.250   1.1472   0.01560   0.00896  -0.0815   0.2966   1.0000
   7.500   1.1646   0.01599   0.00930  -0.0797   0.2896   1.0000
   7.750   1.1845   0.01628   0.00959  -0.0784   0.2838   1.0000
   8.000   1.2019   0.01668   0.00995  -0.0767   0.2755   1.0000
   8.250   1.2196   0.01708   0.01032  -0.0750   0.2680   1.0000
   8.500   1.2361   0.01754   0.01073  -0.0733   0.2577   1.0000
   8.750   1.2530   0.01798   0.01115  -0.0716   0.2493   1.0000
   9.000   1.2679   0.01853   0.01164  -0.0696   0.2391   1.0000
   9.250   1.2835   0.01906   0.01213  -0.0678   0.2279   1.0000
   9.500   1.2975   0.01967   0.01269  -0.0659   0.2170   1.0000
   9.750   1.3097   0.02038   0.01334  -0.0637   0.2052   1.0000
  10.000   1.3219   0.02112   0.01402  -0.0616   0.1944   1.0000
  10.250   1.3347   0.02183   0.01472  -0.0596   0.1872   1.0000
  10.500   1.3471   0.02259   0.01544  -0.0576   0.1810   1.0000
  10.750   1.3594   0.02336   0.01620  -0.0557   0.1757   1.0000
  11.000   1.3716   0.02415   0.01699  -0.0538   0.1710   1.0000
  11.250   1.3838   0.02496   0.01781  -0.0520   0.1673   1.0000
  11.500   1.3933   0.02596   0.01880  -0.0499   0.1629   1.0000
  11.750   1.4069   0.02673   0.01961  -0.0484   0.1606   1.0000
  12.000   1.4193   0.02758   0.02050  -0.0468   0.1579   1.0000
  12.250   1.4309   0.02851   0.02146  -0.0452   0.1556   1.0000
  12.500   1.4411   0.02956   0.02252  -0.0435   0.1531   1.0000
  12.750   1.4495   0.03076   0.02375  -0.0417   0.1504   1.0000
  13.000   1.4590   0.03192   0.02494  -0.0402   0.1483   1.0000
  13.250   1.4711   0.03293   0.02601  -0.0389   0.1464   1.0000
  13.500   1.4815   0.03407   0.02720  -0.0376   0.1444   1.0000
  13.750   1.4912   0.03530   0.02848  -0.0362   0.1420   1.0000
  14.000   1.4982   0.03679   0.03001  -0.0348   0.1392   1.0000
  14.250   1.5027   0.03852   0.03176  -0.0334   0.1359   1.0000
  14.500   1.5117   0.03993   0.03323  -0.0323   0.1336   1.0000
  14.750   1.5201   0.04142   0.03478  -0.0313   0.1304   1.0000
  15.000   1.5242   0.04333   0.03670  -0.0302   0.1254   1.0000
  15.250   1.5281   0.04531   0.03872  -0.0292   0.1218   1.0000
  15.500   1.5320   0.04732   0.04077  -0.0283   0.1155   1.0000
  15.750   1.5325   0.04974   0.04320  -0.0275   0.1096   1.0000
  16.000   1.5297   0.05255   0.04601  -0.0267   0.1014   1.0000
  16.250   1.5261   0.05549   0.04896  -0.0260   0.0949   1.0000
  16.500   1.5207   0.05870   0.05219  -0.0254   0.0901   1.0000
  16.750   1.5174   0.06175   0.05529  -0.0250   0.0873   1.0000
  17.000   1.5114   0.06515   0.05873  -0.0247   0.0842   1.0000
  17.250   1.5072   0.06839   0.06203  -0.0245   0.0823   1.0000
  17.500   1.5049   0.07143   0.06515  -0.0244   0.0808   1.0000
  17.750   1.5004   0.07481   0.06859  -0.0244   0.0792   1.0000
  18.000   1.4937   0.07849   0.07233  -0.0246   0.0774   1.0000
  18.250   1.4878   0.08211   0.07602  -0.0248   0.0763   1.0000
  18.500   1.4815   0.08580   0.07978  -0.0251   0.0749   1.0000
  18.750   1.4784   0.08915   0.08320  -0.0255   0.0738   1.0000
<< Back to GOE 481A AIRFOIL (goe481a-il)

Polar data table (+)

Polar graphs


<< Back to GOE 481A AIRFOIL (goe481a-il)