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

GOE 63 AIRFOIL (goe63-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 63 AIRFOIL (goe63-il)
Reynolds number: 500,000
Max Cl/Cd: 110.32 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe63-il-500000.txt
Download as CSV file: xf-goe63-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 63 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.1828   0.09403   0.09155  -0.0367   0.8703   0.0215
  -7.750  -0.1745   0.09157   0.08898  -0.0371   0.8470   0.0217
  -7.500  -0.1655   0.08916   0.08647  -0.0380   0.8295   0.0219
  -7.250  -0.1565   0.08680   0.08405  -0.0392   0.8152   0.0221
  -7.000  -0.1444   0.08427   0.08144  -0.0413   0.8034   0.0224
  -6.750  -0.1298   0.08155   0.07869  -0.0441   0.7925   0.0228
  -6.500  -0.1138   0.07878   0.07586  -0.0473   0.7831   0.0233
  -6.250  -0.0952   0.07587   0.07290  -0.0511   0.7737   0.0240
  -6.000  -0.0662   0.07268   0.06962  -0.0595   0.7653   0.0245
  -5.750  -0.0305   0.06910   0.06593  -0.0703   0.7570   0.0246
  -5.500  -0.0118   0.06567   0.06244  -0.0722   0.7505   0.0247
  -5.250   0.0028   0.06279   0.05956  -0.0719   0.7434   0.0249
  -5.000   0.0221   0.06016   0.05688  -0.0734   0.7365   0.0251
  -4.750   0.0453   0.05758   0.05426  -0.0760   0.7294   0.0253
  -4.500   0.0709   0.05505   0.05167  -0.0791   0.7221   0.0257
  -4.250   0.0989   0.05249   0.04902  -0.0827   0.7157   0.0264
  -4.000   0.1295   0.04977   0.04624  -0.0866   0.7090   0.0271
  -3.750   0.1852   0.04636   0.04257  -0.0964   0.7029   0.0284
  -3.500   0.2187   0.04304   0.03911  -0.1002   0.6974   0.0286
  -3.250   0.2399   0.04069   0.03677  -0.1010   0.6913   0.0288
  -3.000   0.2655   0.03868   0.03470  -0.1024   0.6854   0.0291
  -2.750   0.2948   0.03675   0.03268  -0.1044   0.6799   0.0296
  -2.500   0.3266   0.03475   0.03061  -0.1067   0.6735   0.0304
  -2.250   0.3606   0.03278   0.02848  -0.1090   0.6676   0.0317
  -2.000   0.4102   0.03008   0.02540  -0.1129   0.6626   0.0330
  -1.750   0.4361   0.02802   0.02335  -0.1143   0.6576   0.0334
  -1.500   0.4643   0.02661   0.02187  -0.1154   0.6526   0.0339
  -1.250   0.4946   0.02529   0.02043  -0.1165   0.6477   0.0348
  -1.000   0.5360   0.02404   0.01882  -0.1176   0.6426   0.0380
  -0.750   0.5656   0.02180   0.01653  -0.1191   0.6373   0.0386
  -0.500   0.5943   0.02079   0.01544  -0.1199   0.6324   0.0394
  -0.250   0.6246   0.01984   0.01441  -0.1207   0.6273   0.0408
   0.000   0.6598   0.01860   0.01287  -0.1212   0.6216   0.0447
   0.250   0.6887   0.01767   0.01188  -0.1219   0.6158   0.0457
   0.500   0.7185   0.01695   0.01110  -0.1224   0.6100   0.0474
   0.750   0.7511   0.01619   0.01006  -0.1226   0.6038   0.0522
   1.000   0.7799   0.01543   0.00926  -0.1231   0.5973   0.0538
   1.250   0.8096   0.01580   0.00949  -0.1228   0.5898   0.0599
   1.500   0.8397   0.01427   0.00790  -0.1236   0.5823   0.0630
   1.750   0.8685   0.01480   0.00832  -0.1232   0.5729   0.0702
   2.250   0.9310   0.01184   0.00489  -0.1234   0.5532   0.0446
   2.500   0.9596   0.01156   0.00456  -0.1233   0.5414   0.0439
   2.750   0.9880   0.01130   0.00424  -0.1233   0.5274   0.0441
   3.000   1.0165   0.01097   0.00389  -0.1234   0.5102   0.0451
   3.250   1.0443   0.01094   0.00380  -0.1234   0.4870   0.0470
   3.500   1.0713   0.01107   0.00382  -0.1232   0.4591   0.0500
   3.750   1.0979   0.01125   0.00384  -0.1230   0.4277   0.0522
   4.000   1.1241   0.01152   0.00397  -0.1228   0.3995   0.0570
   4.250   1.1506   0.01177   0.00414  -0.1226   0.3779   0.0689
   4.500   1.1705   0.01061   0.00446  -0.1213   0.3630   1.0000
   4.750   1.1966   0.01095   0.00467  -0.1210   0.3492   1.0000
   5.000   1.2227   0.01127   0.00490  -0.1208   0.3376   1.0000
   5.250   1.2491   0.01155   0.00513  -0.1206   0.3283   1.0000
   5.750   1.3012   0.01214   0.00561  -0.1201   0.3105   1.0000
   6.250   1.3528   0.01273   0.00614  -0.1195   0.2947   1.0000
   6.500   1.3779   0.01308   0.00643  -0.1191   0.2868   1.0000
   6.750   1.4038   0.01332   0.00670  -0.1189   0.2794   1.0000
   7.000   1.4285   0.01367   0.00700  -0.1185   0.2709   1.0000
   7.250   1.4540   0.01393   0.00727  -0.1182   0.2626   1.0000
   7.750   1.5029   0.01460   0.00791  -0.1173   0.2450   1.0000
   8.000   1.5264   0.01499   0.00827  -0.1167   0.2354   1.0000
   8.250   1.5496   0.01539   0.00863  -0.1161   0.2245   1.0000
   8.500   1.5725   0.01580   0.00902  -0.1155   0.2132   1.0000
   8.750   1.5944   0.01628   0.00946  -0.1147   0.2020   1.0000
   9.000   1.6150   0.01683   0.00995  -0.1137   0.1913   1.0000
   9.250   1.6357   0.01734   0.01044  -0.1128   0.1817   1.0000
   9.500   1.6557   0.01787   0.01098  -0.1117   0.1735   1.0000
   9.750   1.6739   0.01850   0.01157  -0.1104   0.1656   1.0000
  10.000   1.6930   0.01903   0.01213  -0.1092   0.1590   1.0000
  10.250   1.7093   0.01969   0.01277  -0.1076   0.1526   1.0000
  10.500   1.7255   0.02031   0.01342  -0.1060   0.1473   1.0000
  10.750   1.7408   0.02092   0.01406  -0.1042   0.1419   1.0000
  11.000   1.7497   0.02176   0.01490  -0.1015   0.1367   1.0000
  11.250   1.7635   0.02241   0.01561  -0.0996   0.1328   1.0000
  11.500   1.7755   0.02321   0.01646  -0.0976   0.1283   1.0000
  11.750   1.7838   0.02431   0.01756  -0.0954   0.1236   1.0000
  12.000   1.7969   0.02515   0.01848  -0.0938   0.1199   1.0000
  12.250   1.8081   0.02617   0.01955  -0.0922   0.1156   1.0000
  12.500   1.8153   0.02757   0.02096  -0.0905   0.1110   1.0000
  12.750   1.8271   0.02867   0.02215  -0.0892   0.1069   1.0000
  13.000   1.8355   0.03011   0.02361  -0.0879   0.1012   1.0000
  13.250   1.8432   0.03168   0.02523  -0.0867   0.0959   1.0000
  13.500   1.8492   0.03346   0.02701  -0.0854   0.0887   1.0000
  13.750   1.8525   0.03554   0.02910  -0.0842   0.0790   1.0000
  14.000   1.8492   0.03835   0.03185  -0.0829   0.0657   1.0000
  14.250   1.8351   0.04234   0.03572  -0.0813   0.0462   1.0000
  14.500   1.8137   0.04731   0.04061  -0.0799   0.0310   1.0000
  14.750   1.8006   0.05151   0.04484  -0.0790   0.0259   1.0000
  15.000   1.7912   0.05543   0.04885  -0.0784   0.0237   1.0000
  15.250   1.7824   0.05940   0.05293  -0.0780   0.0222   1.0000
  15.500   1.7757   0.06325   0.05689  -0.0778   0.0214   1.0000
  15.750   1.7674   0.06744   0.06120  -0.0778   0.0207   1.0000
  16.000   1.7567   0.07208   0.06597  -0.0781   0.0201   1.0000
  16.250   1.7445   0.07705   0.07107  -0.0786   0.0197   1.0000
  16.500   1.7302   0.08252   0.07667  -0.0795   0.0193   1.0000
  16.750   1.7136   0.08849   0.08279  -0.0807   0.0189   1.0000
  17.000   1.6990   0.09429   0.08872  -0.0820   0.0187   1.0000
  17.250   1.6854   0.10007   0.09465  -0.0835   0.0185   1.0000
  17.500   1.6703   0.10623   0.10095  -0.0853   0.0183   1.0000
  17.750   1.6547   0.11255   0.10741  -0.0873   0.0181   1.0000
  18.000   1.6384   0.11910   0.11410  -0.0896   0.0180   1.0000
<< Back to GOE 63 AIRFOIL (goe63-il)

Polar data table (+)

Polar graphs


<< Back to GOE 63 AIRFOIL (goe63-il)