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

GOE 498 AIRFOIL (goe498-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 498 AIRFOIL (goe498-il)
Reynolds number: 200,000
Max Cl/Cd: 75.94 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe498-il-200000.txt
Download as CSV file: xf-goe498-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 498 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
  ------ -------- --------- --------- -------- -------- --------
 -11.500  -0.3086   0.12861   0.12493  -0.0357   0.9998   0.0883
 -11.250  -0.2873   0.12510   0.12141  -0.0394   0.9972   0.0909
 -11.000  -0.5438   0.06343   0.05939  -0.0884   0.9847   0.0704
 -10.750  -0.5192   0.06076   0.05674  -0.0918   0.9814   0.0712
 -10.500  -0.5110   0.05699   0.05294  -0.0966   0.9729   0.0720
 -10.250  -0.5587   0.03987   0.03460  -0.1212   0.9538   0.0717
 -10.000  -0.5387   0.03866   0.03338  -0.1211   0.9446   0.0729
  -9.750  -0.5090   0.03589   0.03032  -0.1249   0.9404   0.0746
  -9.500  -0.4894   0.03337   0.02741  -0.1264   0.9324   0.0760
  -9.250  -0.4617   0.03110   0.02459  -0.1288   0.9264   0.0776
  -9.000  -0.4253   0.02916   0.02266  -0.1315   0.9239   0.0797
  -8.750  -0.3848   0.02791   0.02135  -0.1345   0.9221   0.0822
  -8.500  -0.3665   0.02687   0.02011  -0.1334   0.9119   0.0842
  -8.250  -0.3298   0.02534   0.01825  -0.1358   0.9086   0.0866
  -8.000  -0.2890   0.02407   0.01708  -0.1385   0.9067   0.0895
  -7.750  -0.2468   0.02304   0.01592  -0.1414   0.9050   0.0932
  -7.500  -0.2261   0.02226   0.01497  -0.1403   0.8955   0.0957
  -7.250  -0.1906   0.02129   0.01410  -0.1418   0.8913   0.0991
  -7.000  -0.1512   0.02050   0.01320  -0.1439   0.8883   0.1035
  -6.750  -0.1114   0.01956   0.01226  -0.1462   0.8858   0.1078
  -6.500  -0.0926   0.01923   0.01194  -0.1444   0.8751   0.1114
  -6.250  -0.0568   0.01856   0.01115  -0.1457   0.8705   0.1164
  -6.000  -0.0186   0.01792   0.01057  -0.1476   0.8672   0.1220
  -5.750   0.0019   0.01769   0.01025  -0.1460   0.8570   0.1269
  -5.500   0.0349   0.01710   0.00972  -0.1468   0.8515   0.1332
  -5.250   0.0718   0.01658   0.00912  -0.1482   0.8473   0.1416
  -5.000   0.0912   0.01633   0.00894  -0.1465   0.8368   0.1489
  -4.750   0.1243   0.01580   0.00842  -0.1473   0.8314   0.1606
  -4.500   0.1509   0.01541   0.00808  -0.1469   0.8237   0.1754
  -4.250   0.1786   0.01504   0.00776  -0.1467   0.8158   0.1949
  -4.000   0.2128   0.01474   0.00741  -0.1476   0.8104   0.2171
  -3.750   0.2358   0.01466   0.00739  -0.1466   0.8010   0.2329
  -3.500   0.2672   0.01449   0.00719  -0.1469   0.7942   0.2489
  -3.250   0.2950   0.01440   0.00710  -0.1466   0.7865   0.2630
  -3.000   0.3232   0.01432   0.00697  -0.1464   0.7782   0.2772
  -2.750   0.3560   0.01422   0.00679  -0.1470   0.7719   0.2922
  -2.500   0.3796   0.01417   0.00680  -0.1459   0.7620   0.3049
  -2.250   0.4115   0.01405   0.00662  -0.1463   0.7543   0.3209
  -2.000   0.4359   0.01407   0.00660  -0.1453   0.7438   0.3360
  -1.750   0.4663   0.01400   0.00650  -0.1455   0.7357   0.3521
  -1.500   0.4911   0.01399   0.00654  -0.1447   0.7266   0.3660
  -1.250   0.5195   0.01398   0.00649  -0.1445   0.7187   0.3816
  -1.000   0.5475   0.01405   0.00651  -0.1442   0.7112   0.3975
  -0.750   0.5727   0.01404   0.00657  -0.1435   0.7025   0.4119
  -0.500   0.6030   0.01405   0.00654  -0.1436   0.6956   0.4277
  -0.250   0.6266   0.01412   0.00663  -0.1426   0.6868   0.4424
   0.000   0.6541   0.01411   0.00664  -0.1423   0.6793   0.4565
   0.250   0.6810   0.01415   0.00670  -0.1418   0.6717   0.4709
   0.500   0.7059   0.01417   0.00674  -0.1410   0.6628   0.4866
   0.750   0.7357   0.01421   0.00671  -0.1411   0.6555   0.5045
   1.000   0.7574   0.01422   0.00685  -0.1397   0.6459   0.5214
   1.250   0.7851   0.01422   0.00684  -0.1393   0.6381   0.5415
   1.500   0.8086   0.01427   0.00696  -0.1383   0.6293   0.5628
   1.750   0.8341   0.01428   0.00699  -0.1375   0.6211   0.5875
   2.000   0.8593   0.01430   0.00708  -0.1368   0.6134   0.6134
   2.250   0.8824   0.01428   0.00717  -0.1356   0.6047   0.6422
   2.500   0.9092   0.01426   0.00717  -0.1351   0.5973   0.6731
   2.750   0.9297   0.01421   0.00730  -0.1334   0.5879   0.7074
   3.000   0.9536   0.01410   0.00729  -0.1322   0.5802   0.7580
   3.250   0.9849   0.01380   0.00738  -0.1325   0.5703   1.0000
   3.500   1.0126   0.01397   0.00738  -0.1324   0.5618   1.0000
   3.750   1.0357   0.01417   0.00755  -0.1315   0.5524   1.0000
   4.000   1.0615   0.01436   0.00764  -0.1310   0.5439   1.0000
   4.250   1.0850   0.01460   0.00784  -0.1301   0.5351   1.0000
   4.500   1.1091   0.01481   0.00798  -0.1293   0.5263   1.0000
   4.750   1.1326   0.01507   0.00819  -0.1284   0.5175   1.0000
   5.000   1.1551   0.01529   0.00837  -0.1273   0.5083   1.0000
   5.250   1.1783   0.01556   0.00859  -0.1264   0.4994   1.0000
   5.500   1.1994   0.01580   0.00881  -0.1251   0.4901   1.0000
   5.750   1.2226   0.01610   0.00905  -0.1242   0.4818   1.0000
   6.000   1.2422   0.01636   0.00931  -0.1226   0.4725   1.0000
   6.250   1.2643   0.01666   0.00954  -0.1215   0.4639   1.0000
   6.500   1.2823   0.01693   0.00984  -0.1197   0.4546   1.0000
   6.750   1.3047   0.01726   0.01009  -0.1187   0.4468   1.0000
   7.000   1.3212   0.01756   0.01045  -0.1166   0.4382   1.0000
   7.250   1.3433   0.01790   0.01071  -0.1156   0.4307   1.0000
   7.500   1.3573   0.01822   0.01111  -0.1131   0.4226   1.0000
   7.750   1.3745   0.01856   0.01141  -0.1112   0.4148   1.0000
   8.000   1.3891   0.01894   0.01181  -0.1089   0.4066   1.0000
   8.250   1.4028   0.01931   0.01217  -0.1064   0.3978   1.0000
   8.500   1.4157   0.01974   0.01261  -0.1039   0.3887   1.0000
   8.750   1.4278   0.02018   0.01303  -0.1014   0.3794   1.0000
   9.000   1.4388   0.02068   0.01357  -0.0987   0.3699   1.0000
   9.250   1.4503   0.02124   0.01406  -0.0962   0.3602   1.0000
   9.500   1.4582   0.02185   0.01473  -0.0933   0.3495   1.0000
   9.750   1.4665   0.02255   0.01538  -0.0906   0.3387   1.0000
  10.000   1.4724   0.02333   0.01617  -0.0877   0.3269   1.0000
  10.250   1.4785   0.02419   0.01706  -0.0849   0.3151   1.0000
  10.500   1.4837   0.02518   0.01800  -0.0822   0.3036   1.0000
  10.750   1.4884   0.02625   0.01907  -0.0796   0.2917   1.0000
  11.000   1.4937   0.02737   0.02022  -0.0772   0.2801   1.0000
  11.250   1.4980   0.02862   0.02145  -0.0748   0.2695   1.0000
  11.500   1.5019   0.02996   0.02279  -0.0726   0.2585   1.0000
  11.750   1.5068   0.03134   0.02419  -0.0706   0.2482   1.0000
  12.000   1.5096   0.03291   0.02572  -0.0686   0.2395   1.0000
  12.250   1.5154   0.03437   0.02724  -0.0670   0.2307   1.0000
  12.500   1.5182   0.03610   0.02894  -0.0653   0.2233   1.0000
  12.750   1.5237   0.03771   0.03061  -0.0639   0.2157   1.0000
  13.000   1.5253   0.03966   0.03251  -0.0624   0.2093   1.0000
  13.250   1.5308   0.04141   0.03435  -0.0613   0.2024   1.0000
  13.500   1.5322   0.04353   0.03644  -0.0601   0.1962   1.0000
  13.750   1.5360   0.04553   0.03851  -0.0591   0.1900   1.0000
  14.000   1.5373   0.04780   0.04079  -0.0581   0.1836   1.0000
  14.250   1.5381   0.05018   0.04317  -0.0572   0.1773   1.0000
  14.500   1.5384   0.05270   0.04576  -0.0565   0.1706   1.0000
  14.750   1.5372   0.05540   0.04843  -0.0558   0.1638   1.0000
  15.000   1.5348   0.05839   0.05154  -0.0554   0.1562   1.0000
  15.250   1.5306   0.06161   0.05478  -0.0550   0.1486   1.0000
  15.500   1.5244   0.06524   0.05848  -0.0550   0.1400   1.0000
  15.750   1.5185   0.06904   0.06238  -0.0551   0.1305   1.0000
  16.000   1.5112   0.07310   0.06650  -0.0555   0.1210   1.0000
  16.250   1.5031   0.07738   0.07079  -0.0561   0.1124   1.0000
  16.500   1.4952   0.08175   0.07517  -0.0568   0.1047   1.0000
  16.750   1.4894   0.08588   0.07933  -0.0575   0.0986   1.0000
  17.000   1.4804   0.09046   0.08386  -0.0584   0.0942   1.0000
<< Back to GOE 498 AIRFOIL (goe498-il)

Polar data table (+)

Polar graphs


<< Back to GOE 498 AIRFOIL (goe498-il)