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

GOE 446 AIRFOIL (goe446-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 446 AIRFOIL (goe446-il)
Reynolds number: 100,000
Max Cl/Cd: 50.86 at α=10.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe446-il-100000.txt
Download as CSV file: xf-goe446-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 446 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.1879   0.10751   0.10283  -0.0685   0.9416   0.1027
  -7.250  -0.2132   0.10797   0.10336  -0.0693   0.9299   0.1036
  -7.000  -0.2238   0.10710   0.10245  -0.0790   0.9186   0.1043
  -6.750  -0.2071   0.10138   0.09680  -0.0732   0.9169   0.1053
  -6.500  -0.1873   0.09766   0.09307  -0.0712   0.9141   0.1068
  -6.250  -0.1646   0.09426   0.08965  -0.0732   0.9109   0.1096
  -6.000  -0.1719   0.09295   0.08837  -0.0712   0.9024   0.1115
  -5.750  -0.1575   0.09006   0.08543  -0.0773   0.8966   0.1167
  -5.500  -0.1581   0.08775   0.08298  -0.0880   0.8860   0.1199
  -5.250  -0.1436   0.08395   0.07925  -0.0847   0.8826   0.1213
  -5.000  -0.1173   0.08059   0.07588  -0.0856   0.8797   0.1239
  -4.750  -0.1193   0.07910   0.07440  -0.0835   0.8705   0.1260
  -4.500  -0.0703   0.07434   0.06925  -0.1000   0.8647   0.1363
  -4.250  -0.0671   0.07189   0.06688  -0.0966   0.8568   0.1375
  -4.000  -0.0453   0.06910   0.06410  -0.0966   0.8514   0.1403
  -3.750   0.0100   0.06488   0.05950  -0.1084   0.8479   0.1535
  -3.500   0.0065   0.06355   0.05827  -0.1041   0.8385   0.1551
  -3.250   0.0384   0.06107   0.05575  -0.1061   0.8343   0.1635
  -3.000   0.0856   0.05741   0.05191  -0.1125   0.8317   0.1753
  -2.750   0.0961   0.05663   0.05089  -0.1129   0.8212   0.1882
  -2.500   0.1726   0.04809   0.04102  -0.1244   0.8188   0.1190
  -2.250   0.2069   0.04681   0.04016  -0.1265   0.8160   0.1413
  -2.000   0.2394   0.04239   0.03483  -0.1286   0.8083   0.1095
  -1.750   0.2813   0.04004   0.03199  -0.1313   0.8037   0.1090
  -1.500   0.3301   0.03785   0.02937  -0.1346   0.8010   0.1111
  -1.250   0.3819   0.03602   0.02700  -0.1380   0.7991   0.1187
  -0.750   0.4343   0.03477   0.02563  -0.1378   0.7846   0.1409
  -0.500   0.4567   0.03461   0.02545  -0.1372   0.7765   0.1602
  -0.250   0.4906   0.03403   0.02478  -0.1379   0.7705   0.1850
   0.000   0.5355   0.03294   0.02379  -0.1404   0.7679   0.2153
   0.250   0.5855   0.03163   0.02244  -0.1434   0.7665   0.2357
   0.500   0.5893   0.03248   0.02337  -0.1399   0.7535   0.2432
   0.750   0.6361   0.03125   0.02221  -0.1422   0.7515   0.2632
   1.000   0.6842   0.02995   0.02100  -0.1447   0.7499   0.2943
   1.250   0.6921   0.03065   0.02188  -0.1418   0.7370   0.3286
   1.500   0.7319   0.02818   0.02078  -0.1421   0.7349   1.0000
   1.750   0.7787   0.02738   0.01969  -0.1443   0.7327   1.0000
   2.000   0.7851   0.02851   0.02076  -0.1411   0.7197   1.0000
   2.250   0.8300   0.02778   0.01985  -0.1432   0.7171   1.0000
   2.500   0.8379   0.02887   0.02091  -0.1403   0.7047   1.0000
   2.750   0.8817   0.02814   0.02005  -0.1422   0.7016   1.0000
   3.000   0.9282   0.02732   0.01911  -0.1444   0.6991   1.0000
   3.250   0.9328   0.02855   0.02034  -0.1411   0.6862   1.0000
   3.500   0.9770   0.02787   0.01956  -0.1431   0.6834   1.0000
   3.750   0.9804   0.02928   0.02100  -0.1397   0.6711   1.0000
   4.000   1.0231   0.02862   0.02028  -0.1415   0.6678   1.0000
   4.250   1.0277   0.02999   0.02167  -0.1384   0.6564   1.0000
   4.500   1.0670   0.02952   0.02116  -0.1397   0.6524   1.0000
   4.750   1.1117   0.02887   0.02046  -0.1419   0.6498   1.0000
   5.000   1.1053   0.03086   0.02253  -0.1373   0.6376   1.0000
   5.250   1.1495   0.03015   0.02180  -0.1394   0.6346   1.0000
   5.500   1.1410   0.03237   0.02409  -0.1347   0.6233   1.0000
   5.750   1.1806   0.03190   0.02362  -0.1361   0.6197   1.0000
   6.000   1.2282   0.03107   0.02278  -0.1387   0.6172   1.0000
   6.250   1.2071   0.03395   0.02578  -0.1324   0.6052   1.0000
   6.500   1.2539   0.03307   0.02491  -0.1347   0.6025   1.0000
   6.750   1.2245   0.03646   0.02841  -0.1275   0.5905   1.0000
   7.000   1.2801   0.03491   0.02688  -0.1308   0.5877   1.0000
   7.250   1.2632   0.03742   0.02948  -0.1250   0.5771   1.0000
   7.500   1.3610   0.03236   0.02435  -0.1325   0.5719   1.0000
   7.750   1.3831   0.03210   0.02414  -0.1310   0.5618   1.0000
   8.000   1.4341   0.03059   0.02259  -0.1336   0.5545   1.0000
   8.250   1.4368   0.03177   0.02390  -0.1300   0.5458   1.0000
   8.500   1.4766   0.03114   0.02330  -0.1314   0.5390   1.0000
   8.750   1.4850   0.03199   0.02427  -0.1285   0.5305   1.0000
   9.000   1.5224   0.03145   0.02376  -0.1295   0.5227   1.0000
   9.250   1.5283   0.03236   0.02480  -0.1262   0.5140   1.0000
   9.500   1.5799   0.03119   0.02359  -0.1292   0.5047   1.0000
   9.750   1.5703   0.03239   0.02497  -0.1234   0.4942   1.0000
  10.000   1.5887   0.03225   0.02485  -0.1215   0.4817   1.0000
  10.250   1.6073   0.03201   0.02459  -0.1195   0.4677   1.0000
  10.500   1.6225   0.03199   0.02457  -0.1171   0.4534   1.0000
  10.750   1.6355   0.03216   0.02472  -0.1145   0.4391   1.0000
  11.000   1.6314   0.03312   0.02579  -0.1099   0.4259   1.0000
  11.250   1.6273   0.03420   0.02697  -0.1056   0.4121   1.0000
  11.500   1.6242   0.03535   0.02819  -0.1016   0.3976   1.0000
  11.750   1.6204   0.03667   0.02961  -0.0978   0.3828   1.0000
  12.000   1.6146   0.03827   0.03128  -0.0942   0.3675   1.0000
  12.250   1.6067   0.04019   0.03328  -0.0907   0.3510   1.0000
  12.500   1.5974   0.04240   0.03555  -0.0875   0.3331   1.0000
  12.750   1.5886   0.04476   0.03792  -0.0846   0.3141   1.0000
  13.000   1.5792   0.04740   0.04053  -0.0820   0.2950   1.0000
  13.250   1.5707   0.05023   0.04333  -0.0798   0.2766   1.0000
  13.500   1.5635   0.05308   0.04613  -0.0778   0.2597   1.0000
  13.750   1.5565   0.05605   0.04905  -0.0760   0.2443   1.0000
  14.000   1.5485   0.05927   0.05223  -0.0744   0.2296   1.0000
  14.250   1.5410   0.06258   0.05552  -0.0730   0.2163   1.0000
  14.500   1.5338   0.06595   0.05887  -0.0719   0.2042   1.0000
  14.750   1.5241   0.06974   0.06264  -0.0709   0.1915   1.0000
  15.000   1.5136   0.07386   0.06681  -0.0703   0.1789   1.0000
  15.250   1.5024   0.07824   0.07126  -0.0699   0.1662   1.0000
  15.500   1.4893   0.08300   0.07606  -0.0698   0.1529   1.0000
  15.750   1.4744   0.08817   0.08124  -0.0699   0.1388   1.0000
  16.000   1.4582   0.09366   0.08671  -0.0703   0.1242   1.0000
  16.250   1.4418   0.09926   0.09225  -0.0709   0.1099   1.0000
  16.500   1.4277   0.10463   0.09755  -0.0716   0.0968   1.0000
<< Back to GOE 446 AIRFOIL (goe446-il)

Polar data table (+)

Polar graphs


<< Back to GOE 446 AIRFOIL (goe446-il)