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

GOE 458 AIRFOIL (goe458-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 458 AIRFOIL (goe458-il)
Reynolds number: 500,000
Max Cl/Cd: 117.84 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe458-il-500000.txt
Download as CSV file: xf-goe458-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 458 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3093   0.10206   0.09992  -0.0320   1.0000   0.0099
  -8.500  -0.3136   0.10018   0.09807  -0.0305   1.0000   0.0098
  -8.250  -0.3165   0.09807   0.09601  -0.0296   0.9996   0.0103
  -8.000  -0.2994   0.09391   0.09184  -0.0344   0.9973   0.0109
  -7.750  -0.2810   0.08967   0.08761  -0.0399   0.9943   0.0118
  -7.500  -0.2626   0.08574   0.08368  -0.0462   0.9891   0.0125
  -7.250  -0.2391   0.08170   0.07965  -0.0535   0.9843   0.0129
  -7.000  -0.2169   0.07710   0.07504  -0.0602   0.9787   0.0130
  -6.750  -0.1917   0.07212   0.07005  -0.0681   0.9737   0.0131
  -6.500  -0.1677   0.06387   0.06179  -0.0793   0.9703   0.0136
  -6.250  -0.1472   0.06030   0.05821  -0.0835   0.9651   0.0146
  -6.000  -0.1186   0.05670   0.05455  -0.0895   0.9608   0.0158
  -5.750  -0.0835   0.05172   0.04951  -0.0984   0.9570   0.0167
  -5.500  -0.0537   0.04700   0.04470  -0.1052   0.9491   0.0178
  -5.250  -0.0179   0.04155   0.03912  -0.1129   0.9432   0.0191
  -5.000   0.0142   0.03642   0.03381  -0.1178   0.9340   0.0209
  -4.750   0.0484   0.03142   0.02844  -0.1220   0.9271   0.0218
  -4.500   0.0666   0.02377   0.02019  -0.1248   0.9175   0.0231
  -4.250   0.0892   0.02152   0.01780  -0.1251   0.9098   0.0248
  -4.000   0.1148   0.01974   0.01578  -0.1253   0.9030   0.0273
  -3.750   0.1405   0.01821   0.01391  -0.1249   0.8949   0.0309
  -3.500   0.1679   0.01307   0.00781  -0.1234   0.8884   0.0161
  -3.250   0.1934   0.01185   0.00635  -0.1226   0.8802   0.0168
  -3.000   0.2206   0.01112   0.00545  -0.1223   0.8735   0.0184
  -2.750   0.2467   0.01042   0.00463  -0.1217   0.8651   0.0197
  -2.500   0.2731   0.00967   0.00373  -0.1211   0.8572   0.0201
  -2.250   0.2998   0.00917   0.00311  -0.1206   0.8489   0.0214
  -2.000   0.3260   0.00864   0.00242  -0.1200   0.8400   0.0248
  -1.750   0.3530   0.00833   0.00204  -0.1196   0.8317   0.0371
  -1.500   0.3788   0.00790   0.00190  -0.1192   0.8221   0.1521
  -1.250   0.4053   0.00783   0.00186  -0.1189   0.8119   0.1752
  -1.000   0.4320   0.00779   0.00178  -0.1186   0.8017   0.1920
  -0.750   0.4586   0.00774   0.00170  -0.1183   0.7913   0.2091
  -0.500   0.4850   0.00767   0.00164  -0.1180   0.7800   0.2280
  -0.250   0.5111   0.00759   0.00159  -0.1176   0.7671   0.2544
   0.000   0.5369   0.00749   0.00155  -0.1171   0.7526   0.2997
   0.250   0.5620   0.00732   0.00153  -0.1166   0.7364   0.3777
   0.500   0.6000   0.00600   0.00158  -0.1191   0.7200   1.0000
   0.750   0.6253   0.00610   0.00157  -0.1185   0.7040   1.0000
   1.000   0.6507   0.00622   0.00157  -0.1179   0.6892   1.0000
   1.250   0.6762   0.00635   0.00160  -0.1173   0.6751   1.0000
   1.500   0.7015   0.00649   0.00164  -0.1168   0.6612   1.0000
   1.750   0.7267   0.00663   0.00170  -0.1162   0.6469   1.0000
   2.000   0.7521   0.00678   0.00177  -0.1157   0.6340   1.0000
   2.250   0.7779   0.00691   0.00185  -0.1153   0.6236   1.0000
   2.500   0.8036   0.00706   0.00197  -0.1149   0.6147   1.0000
   2.750   0.8296   0.00719   0.00208  -0.1145   0.6055   1.0000
   3.000   0.8554   0.00732   0.00220  -0.1141   0.5963   1.0000
   3.250   0.8796   0.00749   0.00231  -0.1133   0.5773   1.0000
   3.500   0.9038   0.00767   0.00243  -0.1126   0.5567   1.0000
   3.750   0.9270   0.00789   0.00258  -0.1117   0.5283   1.0000
   4.000   0.9503   0.00812   0.00274  -0.1108   0.4997   1.0000
   4.250   0.9694   0.00863   0.00295  -0.1092   0.4313   1.0000
   4.500   0.9737   0.01051   0.00376  -0.1054   0.2388   1.0000
   5.000   0.9972   0.01337   0.00551  -0.1001   0.0205   1.0000
   5.250   1.0194   0.01383   0.00612  -0.0990   0.0184   1.0000
   5.500   1.0403   0.01440   0.00679  -0.0977   0.0159   1.0000
   5.750   1.0591   0.01515   0.00764  -0.0961   0.0143   1.0000
   6.000   1.0745   0.01620   0.00881  -0.0938   0.0132   1.0000
   6.250   1.0847   0.01763   0.01038  -0.0907   0.0126   1.0000
   6.500   1.0959   0.01900   0.01182  -0.0877   0.0125   1.0000
   6.750   1.1100   0.02022   0.01312  -0.0853   0.0126   1.0000
<< Back to GOE 458 AIRFOIL (goe458-il)

Polar data table (+)

Polar graphs


<< Back to GOE 458 AIRFOIL (goe458-il)