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

GOE 802 B AIRFOIL (goe802b-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 802 B AIRFOIL (goe802b-il)
Reynolds number: 200,000
Max Cl/Cd: 73.36 at α=7°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe802b-il-200000-n5.txt
Download as CSV file: xf-goe802b-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 802 B AIRFOIL                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.1589   0.09258   0.08909  -0.0576   0.9616   0.0277
  -8.250  -0.1449   0.08950   0.08599  -0.0625   0.9502   0.0281
  -8.000  -0.1318   0.08664   0.08311  -0.0681   0.9378   0.0282
  -7.750  -0.1174   0.08284   0.07930  -0.0710   0.9305   0.0285
  -7.500  -0.1070   0.07972   0.07617  -0.0709   0.9228   0.0289
  -7.250  -0.0934   0.07686   0.07330  -0.0729   0.9145   0.0295
  -7.000  -0.0775   0.07392   0.07033  -0.0758   0.9073   0.0301
  -6.750  -0.0613   0.07104   0.06742  -0.0789   0.8983   0.0309
  -6.500  -0.0421   0.06799   0.06431  -0.0825   0.8912   0.0319
  -6.250  -0.0173   0.06485   0.06109  -0.0879   0.8830   0.0332
  -6.000   0.0201   0.06167   0.05769  -0.0971   0.8747   0.0337
  -5.750   0.0306   0.05823   0.05427  -0.0961   0.8691   0.0341
  -5.500   0.0473   0.05559   0.05161  -0.0970   0.8610   0.0348
  -5.250   0.0697   0.05298   0.04892  -0.0991   0.8541   0.0357
  -5.000   0.0947   0.05036   0.04619  -0.1016   0.8473   0.0368
  -4.750   0.1230   0.04777   0.04348  -0.1046   0.8392   0.0387
  -4.500   0.1640   0.04567   0.04097  -0.1091   0.8326   0.0396
  -4.250   0.1833   0.04256   0.03786  -0.1099   0.8249   0.0400
  -4.000   0.2044   0.04018   0.03544  -0.1106   0.8167   0.0408
  -3.750   0.2296   0.03829   0.03343  -0.1116   0.8094   0.0427
  -3.500   0.2655   0.03722   0.03199  -0.1133   0.8000   0.0457
  -3.250   0.2919   0.03493   0.02952  -0.1142   0.7918   0.0460
  -3.000   0.3150   0.03262   0.02717  -0.1148   0.7805   0.0466
  -2.750   0.3410   0.03092   0.02533  -0.1154   0.7703   0.0477
  -2.500   0.3683   0.02949   0.02372  -0.1159   0.7586   0.0495
  -2.250   0.4005   0.02923   0.02305  -0.1159   0.7479   0.0518
  -1.750   0.4524   0.02559   0.01920  -0.1169   0.7252   0.0536
  -1.500   0.4810   0.02519   0.01853  -0.1168   0.7137   0.0577
  -1.250   0.5094   0.02449   0.01759  -0.1169   0.7012   0.0580
  -0.750   0.5666   0.02083   0.01324  -0.1170   0.6769   0.0394
  -0.500   0.5937   0.01994   0.01219  -0.1170   0.6631   0.0390
  -0.250   0.6209   0.01920   0.01124  -0.1169   0.6500   0.0387
   0.000   0.6480   0.01854   0.01036  -0.1168   0.6373   0.0387
   0.250   0.6751   0.01792   0.00954  -0.1166   0.6244   0.0393
   0.500   0.7016   0.01759   0.00911  -0.1165   0.6115   0.0409
   0.750   0.7280   0.01725   0.00862  -0.1163   0.5984   0.0418
   1.000   0.7546   0.01679   0.00802  -0.1161   0.5854   0.0417
   1.250   0.7809   0.01642   0.00751  -0.1159   0.5736   0.0419
   1.500   0.8071   0.01611   0.00708  -0.1157   0.5621   0.0424
   1.750   0.8333   0.01583   0.00672  -0.1155   0.5515   0.0434
   2.000   0.8591   0.01556   0.00636  -0.1152   0.5410   0.0448
   2.250   0.8851   0.01551   0.00629  -0.1150   0.5302   0.0471
   2.750   0.9365   0.01554   0.00627  -0.1144   0.5088   0.0538
   3.000   0.9617   0.01556   0.00623  -0.1141   0.4991   0.0587
   3.250   0.9877   0.01565   0.00633  -0.1139   0.4893   0.0670
   3.500   1.0131   0.01599   0.00656  -0.1135   0.4801   0.0791
   3.750   1.0390   0.01619   0.00671  -0.1132   0.4695   0.0902
   4.000   1.0634   0.01629   0.00674  -0.1128   0.4580   0.0958
   4.250   1.0877   0.01629   0.00674  -0.1124   0.4450   0.0980
   4.500   1.1119   0.01637   0.00684  -0.1120   0.4325   0.0993
   4.750   1.1354   0.01657   0.00699  -0.1114   0.4212   0.1011
   5.000   1.1596   0.01664   0.00713  -0.1110   0.4098   0.1020
   5.250   1.1830   0.01680   0.00732  -0.1105   0.3992   0.1033
   5.750   1.2302   0.01716   0.00775  -0.1096   0.3819   0.1063
   6.000   1.2540   0.01731   0.00798  -0.1092   0.3728   0.1086
   6.250   1.2762   0.01755   0.00821  -0.1086   0.3632   0.1128
   6.500   1.2994   0.01772   0.00840  -0.1081   0.3526   0.1188
   6.750   1.3210   0.01803   0.00866  -0.1073   0.3429   0.1228
   7.000   1.3432   0.01831   0.00896  -0.1066   0.3328   0.1251
   7.250   1.3640   0.01869   0.00933  -0.1056   0.3229   0.1267
   7.500   1.3852   0.01904   0.00970  -0.1048   0.3123   0.1284
   7.750   1.4054   0.01945   0.01010  -0.1038   0.3028   0.1303
   8.000   1.4258   0.01983   0.01054  -0.1028   0.2930   0.1326
   8.250   1.4450   0.02027   0.01102  -0.1017   0.2835   0.1367
   8.500   1.4648   0.02013   0.01174  -0.1011   0.2712   0.7146
   9.000   1.4919   0.02080   0.01277  -0.0966   0.2445   1.0000
   9.250   1.5029   0.02154   0.01345  -0.0943   0.2254   1.0000
   9.500   1.5119   0.02246   0.01427  -0.0918   0.2050   1.0000
   9.750   1.5196   0.02352   0.01525  -0.0893   0.1870   1.0000
  10.000   1.5276   0.02463   0.01630  -0.0871   0.1728   1.0000
  10.250   1.5354   0.02580   0.01745  -0.0849   0.1623   1.0000
  10.500   1.5452   0.02689   0.01855  -0.0830   0.1527   1.0000
  10.750   1.5528   0.02816   0.01983  -0.0811   0.1446   1.0000
  11.000   1.5629   0.02930   0.02104  -0.0795   0.1368   1.0000
  11.250   1.5692   0.03075   0.02250  -0.0777   0.1296   1.0000
  11.500   1.5785   0.03201   0.02384  -0.0762   0.1225   1.0000
  11.750   1.5840   0.03362   0.02546  -0.0746   0.1143   1.0000
  12.000   1.5902   0.03522   0.02710  -0.0731   0.1058   1.0000
  12.250   1.5933   0.03714   0.02903  -0.0716   0.0973   1.0000
  12.500   1.5989   0.03890   0.03087  -0.0703   0.0901   1.0000
  12.750   1.6008   0.04106   0.03305  -0.0691   0.0824   1.0000
  13.000   1.6022   0.04332   0.03535  -0.0679   0.0752   1.0000
  13.250   1.6011   0.04592   0.03797  -0.0668   0.0688   1.0000
  13.500   1.6002   0.04859   0.04069  -0.0659   0.0635   1.0000
  13.750   1.5965   0.05165   0.04378  -0.0651   0.0586   1.0000
  14.000   1.5959   0.05448   0.04670  -0.0645   0.0540   1.0000
  14.250   1.5929   0.05770   0.05001  -0.0641   0.0490   1.0000
  14.500   1.5898   0.06104   0.05343  -0.0639   0.0433   1.0000
  14.750   1.5811   0.06517   0.05759  -0.0638   0.0270   1.0000
  15.000   1.5660   0.07032   0.06271  -0.0641   0.0221   1.0000
  15.250   1.5548   0.07511   0.06757  -0.0646   0.0202   1.0000
  15.500   1.5451   0.07979   0.07236  -0.0652   0.0192   1.0000
  15.750   1.5356   0.08455   0.07726  -0.0660   0.0184   1.0000
  16.000   1.5260   0.08944   0.08229  -0.0669   0.0178   1.0000
  16.250   1.5159   0.09447   0.08746  -0.0680   0.0173   1.0000
<< Back to GOE 802 B AIRFOIL (goe802b-il)

Polar data table (+)

Polar graphs


<< Back to GOE 802 B AIRFOIL (goe802b-il)