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

GOE 482 AIRFOIL (goe482-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 482 AIRFOIL (goe482-il)
Reynolds number: 200,000
Max Cl/Cd: 71.23 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe482-il-200000.txt
Download as CSV file: xf-goe482-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 482 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
  ------ -------- --------- --------- -------- -------- --------
 -11.250   0.2056   0.10264   0.09854  -0.1321   0.8728   0.0631
 -11.000   0.2169   0.10029   0.09612  -0.1334   0.8650   0.0652
 -10.750   0.1949   0.10035   0.09616  -0.1373   0.8550   0.0685
 -10.500   0.2115   0.09620   0.09194  -0.1377   0.8486   0.0691
 -10.250   0.2320   0.09324   0.08895  -0.1368   0.8409   0.0701
 -10.000   0.2469   0.09105   0.08668  -0.1371   0.8342   0.0714
  -9.750   0.2556   0.08919   0.08481  -0.1373   0.8265   0.0738
  -9.500   0.2311   0.08915   0.08476  -0.1410   0.8175   0.0780
  -9.250   0.2376   0.08582   0.08141  -0.1413   0.8107   0.0787
  -9.000   0.2605   0.08295   0.07850  -0.1399   0.8034   0.0795
  -8.750   0.2793   0.08081   0.07627  -0.1395   0.7971   0.0811
  -8.500   0.2881   0.07916   0.07464  -0.1391   0.7888   0.0831
  -8.250   0.2936   0.07741   0.07285  -0.1396   0.7815   0.0862
  -8.000   0.2517   0.07668   0.07222  -0.1436   0.7718   0.0893
  -7.750   0.2813   0.07319   0.06866  -0.1419   0.7656   0.0902
  -7.500   0.3028   0.07114   0.06651  -0.1410   0.7597   0.0914
  -7.250   0.3145   0.06961   0.06500  -0.1399   0.7507   0.0933
  -7.000   0.3233   0.06789   0.06322  -0.1399   0.7439   0.0962
  -6.750   0.2817   0.06637   0.06182  -0.1489   0.7335   0.1015
  -6.500   0.2996   0.06367   0.05906  -0.1444   0.7276   0.1023
  -6.250   0.3167   0.06199   0.05735  -0.1419   0.7204   0.1033
  -6.000   0.3317   0.06048   0.05581  -0.1406   0.7125   0.1049
  -5.750   0.3362   0.05579   0.05099  -0.1590   0.7053   0.1152
  -5.500   0.3472   0.05430   0.04955  -0.1538   0.6972   0.1160
  -5.250   0.3639   0.05302   0.04820  -0.1507   0.6912   0.1170
  -5.000   0.3931   0.04822   0.04310  -0.1687   0.6827   0.1299
  -4.750   0.4072   0.04602   0.04096  -0.1655   0.6763   0.1308
  -4.500   0.4234   0.04493   0.03988  -0.1629   0.6699   0.1318
  -4.250   0.4399   0.04393   0.03888  -0.1610   0.6623   0.1333
  -4.000   0.5016   0.02783   0.02133  -0.1878   0.6578   0.1029
  -3.750   0.5299   0.02595   0.01926  -0.1889   0.6512   0.0984
  -3.500   0.5607   0.02375   0.01670  -0.1908   0.6444   0.0957
  -3.250   0.5930   0.02233   0.01486  -0.1923   0.6391   0.0957
  -3.000   0.6210   0.02129   0.01360  -0.1928   0.6321   0.0956
  -2.750   0.6508   0.02038   0.01242  -0.1933   0.6260   0.0955
  -2.500   0.6813   0.01973   0.01149  -0.1939   0.6209   0.0961
  -2.250   0.7080   0.01924   0.01090  -0.1938   0.6141   0.0970
  -2.000   0.7369   0.01883   0.01033  -0.1940   0.6086   0.0985
  -1.750   0.7667   0.01860   0.00989  -0.1943   0.6036   0.1005
  -1.500   0.7929   0.01824   0.00951  -0.1941   0.5975   0.1023
  -1.250   0.8209   0.01787   0.00913  -0.1941   0.5922   0.1045
  -1.000   0.8511   0.01772   0.00887  -0.1946   0.5876   0.1075
  -0.750   0.8764   0.01762   0.00883  -0.1942   0.5819   0.1108
  -0.500   0.9040   0.01755   0.00871  -0.1941   0.5767   0.1152
  -0.250   0.9334   0.01743   0.00858  -0.1945   0.5723   0.1226
   0.000   0.9610   0.01743   0.00858  -0.1945   0.5675   0.1319
   0.250   0.9876   0.01736   0.00861  -0.1944   0.5621   0.1503
   0.500   1.0173   0.01697   0.00861  -0.1951   0.5573   0.2666
   0.750   1.0475   0.01715   0.00885  -0.1956   0.5532   0.3811
   1.000   1.0712   0.01733   0.00914  -0.1950   0.5483   0.4193
   1.250   1.0969   0.01745   0.00933  -0.1946   0.5434   0.4513
   1.500   1.1244   0.01756   0.00947  -0.1946   0.5390   0.4867
   1.750   1.1523   0.01778   0.00970  -0.1947   0.5348   0.5245
   2.000   1.1747   0.01796   0.01002  -0.1938   0.5297   0.5598
   2.250   1.2000   0.01809   0.01023  -0.1934   0.5250   0.6041
   2.500   1.2267   0.01814   0.01039  -0.1932   0.5210   0.6631
   2.750   1.2496   0.01808   0.01059  -0.1921   0.5172   0.8004
   3.000   1.2714   0.01816   0.01080  -0.1911   0.5123   1.0000
   3.250   1.2975   0.01844   0.01099  -0.1909   0.5076   1.0000
   3.500   1.3262   0.01870   0.01110  -0.1912   0.5035   1.0000
   3.750   1.3528   0.01909   0.01140  -0.1912   0.4993   1.0000
   4.000   1.3737   0.01944   0.01178  -0.1903   0.4944   1.0000
   4.250   1.3987   0.01974   0.01203  -0.1899   0.4900   1.0000
   4.500   1.4268   0.02003   0.01220  -0.1902   0.4862   1.0000
   4.750   1.4535   0.02046   0.01256  -0.1902   0.4824   1.0000
   5.000   1.4719   0.02086   0.01304  -0.1888   0.4778   1.0000
   5.250   1.4948   0.02119   0.01335  -0.1882   0.4734   1.0000
   5.500   1.5218   0.02146   0.01354  -0.1883   0.4695   1.0000
   5.750   1.5497   0.02187   0.01385  -0.1885   0.4658   1.0000
   6.000   1.5644   0.02233   0.01443  -0.1866   0.4612   1.0000
   6.250   1.5851   0.02272   0.01484  -0.1857   0.4570   1.0000
   6.500   1.6103   0.02304   0.01511  -0.1855   0.4534   1.0000
   6.750   1.6405   0.02340   0.01536  -0.1862   0.4502   1.0000
   7.000   1.6548   0.02399   0.01606  -0.1843   0.4463   1.0000
   7.250   1.6701   0.02450   0.01666  -0.1826   0.4422   1.0000
   7.500   1.6910   0.02488   0.01704  -0.1817   0.4385   1.0000
   7.750   1.7179   0.02518   0.01727  -0.1819   0.4351   1.0000
   8.000   1.7435   0.02567   0.01774  -0.1819   0.4318   1.0000
   8.250   1.7494   0.02638   0.01861  -0.1787   0.4280   1.0000
   8.500   1.7617   0.02695   0.01925  -0.1766   0.4244   1.0000
   8.750   1.7811   0.02739   0.01971  -0.1756   0.4212   1.0000
   9.000   1.8093   0.02774   0.02001  -0.1760   0.4184   1.0000
   9.250   1.8357   0.02829   0.02055  -0.1763   0.4154   1.0000
   9.500   1.8276   0.02924   0.02168  -0.1711   0.4117   1.0000
   9.750   1.8326   0.02999   0.02253  -0.1681   0.4080   1.0000
  10.000   1.8499   0.03040   0.02295  -0.1669   0.4044   1.0000
  10.250   1.8861   0.03032   0.02272  -0.1683   0.4002   1.0000
  10.500   1.8701   0.03166   0.02427  -0.1627   0.3960   1.0000
  10.750   1.8668   0.03270   0.02542  -0.1590   0.3912   1.0000
  11.000   1.8851   0.03284   0.02549  -0.1580   0.3862   1.0000
  11.250   1.8964   0.03359   0.02626  -0.1564   0.3818   1.0000
  11.500   1.8846   0.03541   0.02826  -0.1524   0.3775   1.0000
  11.750   1.8902   0.03650   0.02942  -0.1505   0.3733   1.0000
  12.000   1.9117   0.03674   0.02961  -0.1500   0.3692   1.0000
  12.250   1.9128   0.03825   0.03121  -0.1478   0.3649   1.0000
  12.500   1.9011   0.04057   0.03371  -0.1447   0.3605   1.0000
  12.750   1.9072   0.04175   0.03492  -0.1432   0.3557   1.0000
  13.000   1.9345   0.04165   0.03470  -0.1432   0.3511   1.0000
  13.250   1.9115   0.04511   0.03842  -0.1399   0.3470   1.0000
  13.500   1.9053   0.04751   0.04095  -0.1379   0.3424   1.0000
  13.750   1.9185   0.04835   0.04178  -0.1371   0.3377   1.0000
  14.000   1.9171   0.05054   0.04405  -0.1356   0.3329   1.0000
  14.250   1.8994   0.05436   0.04806  -0.1337   0.3279   1.0000
  14.500   1.9041   0.05609   0.04981  -0.1328   0.3227   1.0000
  14.750   1.9082   0.05802   0.05178  -0.1319   0.3176   1.0000
  15.000   1.8857   0.06296   0.05693  -0.1306   0.3125   1.0000
  15.250   1.8882   0.06517   0.05918  -0.1299   0.3068   1.0000
  15.500   1.8817   0.06857   0.06266  -0.1293   0.3009   1.0000
  15.750   1.8588   0.07422   0.06850  -0.1288   0.2949   1.0000
  16.000   1.8716   0.07528   0.06948  -0.1285   0.2879   1.0000
  16.250   1.8376   0.08287   0.07734  -0.1286   0.2821   1.0000
  16.500   1.8376   0.08582   0.08029  -0.1286   0.2748   1.0000
  16.750   1.8129   0.09250   0.08716  -0.1291   0.2683   1.0000
  17.000   1.8066   0.09652   0.09118  -0.1295   0.2594   1.0000
  17.250   1.7790   0.10397   0.09884  -0.1306   0.2526   1.0000
  17.500   1.7764   0.10754   0.10238  -0.1311   0.2430   1.0000
<< Back to GOE 482 AIRFOIL (goe482-il)

Polar data table (+)

Polar graphs


<< Back to GOE 482 AIRFOIL (goe482-il)