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

GOE 198 (L.F.G. 5294) AIRFOIL (goe198-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: GOE 198 (L.F.G. 5294) AIRFOIL (goe198-il)
Reynolds number: 1,000,000
Max Cl/Cd: 92.85 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe198-il-1000000-n5.txt
Download as CSV file: xf-goe198-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 198 (L.F.G. 5294) AIRFOIL                   
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.2262   0.09938   0.09713  -0.0438   0.8487   0.0127
  -8.750  -0.2178   0.09643   0.09418  -0.0454   0.8454   0.0130
  -8.500  -0.2169   0.09110   0.08884  -0.0486   0.8413   0.0139
  -8.250  -0.2060   0.08904   0.08677  -0.0497   0.8377   0.0141
  -8.000  -0.1950   0.08699   0.08471  -0.0509   0.8343   0.0142
  -7.750  -0.1842   0.08489   0.08262  -0.0523   0.8304   0.0144
  -7.500  -0.1718   0.08255   0.08027  -0.0545   0.8256   0.0146
  -7.250  -0.1575   0.07973   0.07742  -0.0575   0.8205   0.0149
  -7.000  -0.1415   0.07671   0.07440  -0.0611   0.8155   0.0151
  -6.750  -0.1240   0.07356   0.07121  -0.0650   0.8093   0.0154
  -6.500  -0.1049   0.07029   0.06790  -0.0692   0.8031   0.0156
  -6.250  -0.0799   0.06461   0.06217  -0.0777   0.7970   0.0168
  -6.000  -0.0585   0.06247   0.06000  -0.0804   0.7909   0.0170
  -5.750  -0.0350   0.06004   0.05753  -0.0838   0.7860   0.0172
  -5.500  -0.0096   0.05738   0.05484  -0.0877   0.7803   0.0174
  -5.250   0.0173   0.05453   0.05193  -0.0919   0.7738   0.0177
  -5.000   0.0461   0.05153   0.04886  -0.0964   0.7675   0.0179
  -4.750   0.0766   0.04840   0.04566  -0.1011   0.7594   0.0181
  -4.500   0.1085   0.04525   0.04242  -0.1057   0.7517   0.0182
  -4.250   0.1418   0.04215   0.03922  -0.1103   0.7426   0.0188
  -4.000   0.1778   0.03857   0.03551  -0.1153   0.7344   0.0191
  -3.750   0.2190   0.03404   0.03078  -0.1213   0.7253   0.0196
  -3.500   0.2558   0.03056   0.02713  -0.1253   0.7150   0.0197
  -3.250   0.2924   0.02708   0.02342  -0.1290   0.7020   0.0201
  -3.000   0.3269   0.02432   0.02043  -0.1317   0.6848   0.0202
  -2.750   0.3592   0.02238   0.01825  -0.1335   0.6608   0.0204
  -2.500   0.3911   0.02065   0.01626  -0.1350   0.6321   0.0206
  -2.250   0.4233   0.01887   0.01419  -0.1364   0.6053   0.0210
  -2.000   0.4658   0.01239   0.00675  -0.1401   0.5871   0.0218
  -1.750   0.4955   0.01165   0.00573  -0.1406   0.5694   0.0222
  -1.500   0.5249   0.01112   0.00499  -0.1410   0.5560   0.0225
  -1.250   0.5539   0.01075   0.00443  -0.1413   0.5396   0.0227
  -1.000   0.5826   0.01052   0.00404  -0.1415   0.5228   0.0230
  -0.750   0.6113   0.01035   0.00375  -0.1417   0.5110   0.0232
  -0.500   0.6401   0.01021   0.00354  -0.1419   0.5014   0.0234
  -0.250   0.6687   0.01013   0.00339  -0.1421   0.4910   0.0236
   0.000   0.6974   0.00988   0.00304  -0.1423   0.4774   0.0241
   0.250   0.7258   0.00983   0.00292  -0.1425   0.4614   0.0244
   0.500   0.7542   0.00982   0.00284  -0.1426   0.4441   0.0248
   0.750   0.7820   0.00989   0.00280  -0.1427   0.4171   0.0252
   1.000   0.8081   0.01022   0.00287  -0.1425   0.3623   0.0255
   1.250   0.8349   0.01046   0.00294  -0.1424   0.3316   0.0260
   1.750   0.8904   0.01064   0.00302  -0.1425   0.3087   0.0273
   2.000   0.9181   0.01074   0.00307  -0.1426   0.2990   0.0278
   2.250   0.9459   0.01082   0.00311  -0.1426   0.2889   0.0284
   2.500   0.9735   0.01091   0.00317  -0.1426   0.2789   0.0296
   2.750   1.0008   0.01106   0.00326  -0.1426   0.2670   0.0307
   3.000   1.0278   0.01123   0.00336  -0.1425   0.2521   0.0319
   3.250   1.0538   0.01153   0.00353  -0.1423   0.2242   0.0335
   3.750   1.0946   0.01353   0.00476  -0.1403   0.0586   0.0686
   4.000   1.1210   0.01372   0.00495  -0.1401   0.0548   0.0740
   4.250   1.1474   0.01390   0.00515  -0.1399   0.0535   0.0795
   4.500   1.1736   0.01408   0.00534  -0.1397   0.0526   0.0836
   4.750   1.1997   0.01427   0.00556  -0.1395   0.0517   0.0907
   5.250   1.2476   0.01344   0.00627  -0.1386   0.0501   1.0000
   5.500   1.2730   0.01371   0.00653  -0.1383   0.0494   1.0000
   5.750   1.2981   0.01400   0.00681  -0.1379   0.0487   1.0000
   6.000   1.3230   0.01429   0.00710  -0.1375   0.0480   1.0000
   6.250   1.3475   0.01461   0.00741  -0.1370   0.0474   1.0000
   6.500   1.3714   0.01496   0.00777  -0.1364   0.0463   1.0000
   6.750   1.3955   0.01528   0.00809  -0.1359   0.0454   1.0000
   7.000   1.4197   0.01556   0.00837  -0.1354   0.0450   1.0000
   7.250   1.4435   0.01585   0.00868  -0.1348   0.0445   1.0000
   7.500   1.4669   0.01617   0.00901  -0.1342   0.0440   1.0000
   7.750   1.4900   0.01648   0.00934  -0.1336   0.0432   1.0000
   8.000   1.5127   0.01681   0.00968  -0.1328   0.0424   1.0000
   8.250   1.5348   0.01716   0.01004  -0.1320   0.0416   1.0000
   8.500   1.5561   0.01753   0.01043  -0.1311   0.0406   1.0000
   8.750   1.5764   0.01795   0.01084  -0.1300   0.0392   1.0000
   9.000   1.5963   0.01836   0.01126  -0.1288   0.0381   1.0000
   9.250   1.6168   0.01870   0.01163  -0.1277   0.0373   1.0000
   9.500   1.6364   0.01907   0.01202  -0.1265   0.0359   1.0000
   9.750   1.6543   0.01948   0.01244  -0.1250   0.0338   1.0000
  10.000   1.6696   0.01996   0.01292  -0.1230   0.0314   1.0000
  10.250   1.6788   0.02085   0.01367  -0.1203   0.0174   1.0000
  10.500   1.6893   0.02174   0.01459  -0.1178   0.0145   1.0000
  10.750   1.7011   0.02259   0.01548  -0.1156   0.0132   1.0000
  11.000   1.7124   0.02351   0.01644  -0.1135   0.0121   1.0000
  11.250   1.7251   0.02438   0.01736  -0.1117   0.0116   1.0000
  11.500   1.7372   0.02532   0.01836  -0.1100   0.0111   1.0000
  11.750   1.7483   0.02635   0.01944  -0.1081   0.0106   1.0000
  12.000   1.7584   0.02749   0.02063  -0.1063   0.0101   1.0000
  12.250   1.7672   0.02876   0.02196  -0.1044   0.0097   1.0000
  12.500   1.7747   0.03016   0.02342  -0.1025   0.0092   1.0000
  12.750   1.7833   0.03149   0.02483  -0.1007   0.0090   1.0000
  13.000   1.7917   0.03288   0.02628  -0.0991   0.0087   1.0000
  13.250   1.7991   0.03437   0.02784  -0.0975   0.0085   1.0000
  13.500   1.8054   0.03600   0.02953  -0.0958   0.0082   1.0000
  13.750   1.8107   0.03776   0.03137  -0.0942   0.0079   1.0000
  14.000   1.8148   0.03969   0.03337  -0.0927   0.0077   1.0000
  14.250   1.8175   0.04179   0.03555  -0.0912   0.0075   1.0000
  14.500   1.8188   0.04410   0.03794  -0.0898   0.0073   1.0000
  14.750   1.8188   0.04664   0.04056  -0.0884   0.0071   1.0000
  15.000   1.8168   0.04953   0.04354  -0.0873   0.0069   1.0000
  15.250   1.8157   0.05241   0.04652  -0.0863   0.0068   1.0000
  15.500   1.8155   0.05530   0.04950  -0.0856   0.0067   1.0000
  15.750   1.8142   0.05845   0.05275  -0.0851   0.0066   1.0000
  16.000   1.8120   0.06180   0.05620  -0.0848   0.0065   1.0000
  16.250   1.8084   0.06541   0.05992  -0.0846   0.0064   1.0000
  16.500   1.8037   0.06928   0.06389  -0.0846   0.0063   1.0000
  16.750   1.7982   0.07333   0.06805  -0.0848   0.0062   1.0000
  17.000   1.7915   0.07764   0.07246  -0.0851   0.0061   1.0000
  17.250   1.7839   0.08215   0.07708  -0.0856   0.0060   1.0000
  17.500   1.7751   0.08683   0.08187  -0.0862   0.0060   1.0000
  17.750   1.7658   0.09166   0.08680  -0.0870   0.0059   1.0000
  18.000   1.7557   0.09665   0.09191  -0.0879   0.0058   1.0000
  18.250   1.7453   0.10178   0.09714  -0.0890   0.0058   1.0000
  18.500   1.7345   0.10700   0.10247  -0.0903   0.0057   1.0000
  18.750   1.7234   0.11234   0.10791  -0.0917   0.0056   1.0000
<< Back to GOE 198 (L.F.G. 5294) AIRFOIL (goe198-il)

Polar data table (+)

Polar graphs


<< Back to GOE 198 (L.F.G. 5294) AIRFOIL (goe198-il)