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

GOE 464 AIRFOIL (goe464-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 464 AIRFOIL (goe464-il)
Reynolds number: 1,000,000
Max Cl/Cd: 118.37 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe464-il-1000000.txt
Download as CSV file: xf-goe464-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 464 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250   0.0295   0.09243   0.09033  -0.0816   0.8786   0.0131
  -9.000   0.0412   0.09050   0.08834  -0.0825   0.8664   0.0135
  -8.750   0.0512   0.08844   0.08622  -0.0832   0.8559   0.0139
  -8.500   0.0605   0.08633   0.08408  -0.0838   0.8479   0.0143
  -8.250   0.0691   0.08413   0.08187  -0.0844   0.8400   0.0149
  -8.000   0.0726   0.08158   0.07929  -0.0852   0.8327   0.0154
  -7.750   0.0747   0.07949   0.07720  -0.0853   0.8250   0.0154
  -7.500   0.0815   0.07702   0.07471  -0.0852   0.8189   0.0156
  -7.250   0.0936   0.07522   0.07290  -0.0856   0.8130   0.0158
  -7.000   0.1064   0.07362   0.07130  -0.0861   0.8062   0.0161
  -6.750   0.1191   0.07182   0.06947  -0.0871   0.7994   0.0164
  -6.500   0.1326   0.06979   0.06744  -0.0884   0.7925   0.0167
  -6.250   0.1467   0.06774   0.06536  -0.0898   0.7845   0.0175
  -6.000   0.1640   0.06450   0.06210  -0.0941   0.7760   0.0187
  -5.750   0.1820   0.06188   0.05943  -0.0969   0.7664   0.0187
  -5.500   0.1981   0.05881   0.05635  -0.0990   0.7544   0.0190
  -5.250   0.2152   0.05755   0.05504  -0.0991   0.7383   0.0193
  -5.000   0.2340   0.05581   0.05321  -0.1005   0.7129   0.0197
  -4.750   0.2527   0.05416   0.05140  -0.1018   0.6764   0.0205
  -4.500   0.2840   0.05046   0.04753  -0.1082   0.6533   0.0224
  -4.250   0.3102   0.04680   0.04376  -0.1125   0.6381   0.0227
  -4.000   0.3307   0.04577   0.04266  -0.1128   0.6265   0.0231
  -3.750   0.3539   0.04469   0.04152  -0.1138   0.6165   0.0238
  -3.250   0.4227   0.03842   0.03506  -0.1234   0.6007   0.0274
  -3.000   0.4483   0.03740   0.03400  -0.1244   0.5930   0.0280
  -2.750   0.4773   0.03597   0.03252  -0.1263   0.5853   0.0297
  -2.500   0.5629   0.01835   0.01393  -0.1485   0.5817   0.0302
  -2.250   0.5923   0.01681   0.01219  -0.1494   0.5752   0.0308
  -2.000   0.6205   0.01583   0.01106  -0.1497   0.5677   0.0313
  -1.750   0.6482   0.01514   0.01023  -0.1497   0.5600   0.0321
  -1.500   0.6758   0.01492   0.00996  -0.1495   0.5514   0.0330
  -1.250   0.7028   0.01479   0.00977  -0.1491   0.5421   0.0338
  -1.000   0.7300   0.01473   0.00968  -0.1487   0.5321   0.0345
  -0.750   0.7568   0.01475   0.00966  -0.1483   0.5221   0.0353
  -0.500   0.7834   0.01466   0.00950  -0.1479   0.5106   0.0363
  -0.250   0.8103   0.01454   0.00933  -0.1475   0.4988   0.0373
   0.000   0.8366   0.01458   0.00934  -0.1470   0.4862   0.0398
   0.500   0.8871   0.01538   0.01021  -0.1457   0.4595   0.0520
   0.750   0.9148   0.01468   0.00936  -0.1456   0.4488   0.0636
   1.000   0.9419   0.01440   0.00894  -0.1452   0.4387   0.0686
   1.250   0.9661   0.01506   0.00964  -0.1445   0.4278   0.0697
   1.500   0.9938   0.01510   0.00957  -0.1441   0.4171   0.0752
   1.750   1.0186   0.01634   0.01084  -0.1433   0.4056   0.0813
   2.000   1.0454   0.01615   0.01052  -0.1429   0.3954   0.0815
   2.250   1.0720   0.01589   0.01016  -0.1425   0.3870   0.0816
   2.500   1.0969   0.01493   0.00912  -0.1423   0.3791   0.0823
   2.750   1.1217   0.01473   0.00890  -0.1420   0.3722   0.0828
   3.000   1.1471   0.01456   0.00871  -0.1416   0.3653   0.0834
   3.250   1.1721   0.01446   0.00854  -0.1411   0.3581   0.0842
   3.500   1.1983   0.01430   0.00837  -0.1407   0.3527   0.0855
   3.750   1.2241   0.01422   0.00822  -0.1402   0.3461   0.0871
   4.000   1.2513   0.01451   0.00840  -0.1395   0.3405   0.0889
   4.250   1.2775   0.01438   0.00823  -0.1391   0.3357   0.0890
   4.500   1.3030   0.01426   0.00806  -0.1385   0.3304   0.0890
   4.750   1.3280   0.01417   0.00790  -0.1379   0.3248   0.0891
   5.000   1.3538   0.01401   0.00772  -0.1375   0.3206   0.0892
   5.250   1.3790   0.01388   0.00755  -0.1369   0.3153   0.0892
   5.750   1.4280   0.01324   0.00687  -0.1359   0.3047   0.0907
   6.000   1.4526   0.01324   0.00686  -0.1353   0.2994   0.0913
   6.250   1.4763   0.01332   0.00691  -0.1346   0.2935   0.0923
   6.500   1.5012   0.01332   0.00692  -0.1341   0.2890   0.0940
   6.750   1.5260   0.01366   0.00720  -0.1333   0.2834   0.0972
   7.000   1.5492   0.01380   0.00728  -0.1324   0.2760   0.0973
   7.250   1.5735   0.01383   0.00731  -0.1317   0.2699   0.0973
   7.500   1.5959   0.01362   0.00706  -0.1309   0.2617   0.0984
   7.750   1.6193   0.01368   0.00715  -0.1302   0.2550   0.0992
   8.000   1.6406   0.01392   0.00734  -0.1291   0.2446   0.1004
   8.250   1.6620   0.01416   0.00755  -0.1281   0.2325   0.1027
   8.500   1.6819   0.01473   0.00800  -0.1267   0.2174   0.1059
   8.750   1.7011   0.01482   0.00804  -0.1254   0.2040   0.1079
   9.000   1.7193   0.01520   0.00838  -0.1239   0.1910   0.1094
   9.250   1.7366   0.01565   0.00878  -0.1222   0.1792   0.1126
  10.750   1.8238   0.01808   0.01116  -0.1092   0.1401   0.1123
  11.000   1.8382   0.01848   0.01156  -0.1071   0.1358   0.1078
  11.250   1.8496   0.01907   0.01214  -0.1048   0.1302   0.1074
  11.500   1.8613   0.01965   0.01276  -0.1025   0.1252   0.1089
  11.750   1.8724   0.02029   0.01341  -0.1003   0.1195   0.1094
  12.000   1.8804   0.02115   0.01426  -0.0978   0.1131   0.1097
  12.250   1.8898   0.02196   0.01509  -0.0957   0.1064   0.1101
  12.500   1.8959   0.02303   0.01614  -0.0932   0.0993   0.1101
  12.750   1.8998   0.02430   0.01738  -0.0907   0.0902   0.1108
  13.000   1.8989   0.02599   0.01901  -0.0880   0.0781   0.1111
  13.250   1.8944   0.02805   0.02101  -0.0852   0.0661   0.1113
  13.500   1.8874   0.03044   0.02336  -0.0825   0.0552   0.1121
  13.750   1.8799   0.03304   0.02594  -0.0802   0.0458   0.1130
  14.000   1.8589   0.03709   0.02991  -0.0775   0.0290   0.1132
  14.250   1.8381   0.04153   0.03435  -0.0757   0.0179   0.1132
  14.500   1.8297   0.04503   0.03793  -0.0748   0.0153   0.1135
  14.750   1.8266   0.04815   0.04115  -0.0743   0.0142   0.1139
  15.000   1.8221   0.05159   0.04469  -0.0740   0.0134   0.1144
  15.250   1.8148   0.05554   0.04875  -0.0741   0.0127   0.1148
  15.500   1.8069   0.05971   0.05302  -0.0743   0.0121   0.1152
  15.750   1.8004   0.06380   0.05723  -0.0747   0.0117   0.1153
  16.000   1.7955   0.06777   0.06130  -0.0752   0.0113   0.1171
  16.250   1.7887   0.07213   0.06578  -0.0760   0.0110   0.1175
  16.500   1.7808   0.07678   0.07055  -0.0770   0.0107   0.1186
  16.750   1.7706   0.08191   0.07579  -0.0783   0.0104   0.1193
  17.000   1.7602   0.08722   0.08122  -0.0798   0.0102   0.1200
  17.250   1.7486   0.09287   0.08698  -0.0815   0.0100   0.1206
  17.500   1.7365   0.09871   0.09294  -0.0835   0.0098   0.1212
  17.750   1.7211   0.10526   0.09962  -0.0859   0.0095   0.1213
  18.000   1.7061   0.11186   0.10634  -0.0884   0.0093   0.1222
<< Back to GOE 464 AIRFOIL (goe464-il)

Polar data table (+)

Polar graphs


<< Back to GOE 464 AIRFOIL (goe464-il)