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

GOE 165 (MVA MK.11) AIRFOIL (goe165-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 165 (MVA MK.11) AIRFOIL (goe165-il)
Reynolds number: 200,000
Max Cl/Cd: 69.3 at α=7.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe165-il-200000.txt
Download as CSV file: xf-goe165-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 165 (MVA MK.11) 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3265   0.10288   0.09970  -0.0361   1.0000   0.0487
  -8.500  -0.3353   0.09998   0.09687  -0.0368   1.0000   0.0489
  -8.250  -0.3457   0.09718   0.09415  -0.0367   1.0000   0.0490
  -8.000  -0.3329   0.09289   0.08987  -0.0331   1.0000   0.0507
  -7.750  -0.3314   0.09098   0.08800  -0.0311   1.0000   0.0517
  -7.500  -0.3356   0.08907   0.08616  -0.0292   1.0000   0.0526
  -7.250  -0.3452   0.08731   0.08448  -0.0268   1.0000   0.0535
  -7.000  -0.3576   0.08557   0.08282  -0.0245   1.0000   0.0543
  -6.750  -0.3669   0.08330   0.08063  -0.0236   1.0000   0.0553
  -6.500  -0.3260   0.06192   0.05916  -0.0655   0.9880   0.0621
  -6.250  -0.2988   0.06401   0.06132  -0.0579   0.9850   0.0639
  -6.000  -0.2331   0.04525   0.04205  -0.0974   0.9750   0.0755
  -5.500  -0.1120   0.02854   0.02318  -0.1210   0.9647   0.0366
  -5.250  -0.0651   0.02407   0.01826  -0.1260   0.9614   0.0330
  -5.000  -0.0232   0.02149   0.01528  -0.1287   0.9538   0.0316
  -4.750   0.0210   0.02002   0.01365  -0.1319   0.9481   0.0341
  -4.500   0.0590   0.01855   0.01206  -0.1335   0.9383   0.0350
  -4.250   0.1000   0.01716   0.01061  -0.1357   0.9311   0.0358
  -4.000   0.1394   0.01609   0.00949  -0.1375   0.9220   0.0374
  -3.750   0.1797   0.01470   0.00811  -0.1402   0.9114   0.0417
  -3.500   0.2256   0.01356   0.00679  -0.1438   0.9021   0.0549
  -3.250   0.2754   0.01178   0.00584  -0.1492   0.8923   0.3058
  -3.000   0.3181   0.01204   0.00599  -0.1514   0.8764   0.3466
  -2.750   0.3584   0.01218   0.00597  -0.1533   0.8554   0.3607
  -2.500   0.3963   0.01277   0.00642  -0.1546   0.8342   0.3811
  -2.250   0.4290   0.01306   0.00663  -0.1550   0.8133   0.3887
  -2.000   0.4610   0.01316   0.00646  -0.1556   0.7935   0.3977
  -1.750   0.4902   0.01335   0.00657  -0.1554   0.7754   0.4022
  -1.500   0.5188   0.01350   0.00654  -0.1552   0.7584   0.4087
  -1.250   0.5467   0.01358   0.00646  -0.1550   0.7423   0.4142
  -1.000   0.5737   0.01373   0.00651  -0.1545   0.7274   0.4185
  -0.750   0.6012   0.01387   0.00648  -0.1543   0.7127   0.4256
  -0.500   0.6271   0.01405   0.00663  -0.1536   0.6980   0.4314
  -0.250   0.6533   0.01426   0.00673  -0.1530   0.6823   0.4398
   0.000   0.6786   0.01439   0.00682  -0.1521   0.6653   0.4455
   0.250   0.7039   0.01451   0.00686  -0.1514   0.6467   0.4513
   0.500   0.7293   0.01460   0.00681  -0.1507   0.6260   0.4574
   0.750   0.7530   0.01467   0.00684  -0.1496   0.6004   0.4618
   1.000   0.7764   0.01477   0.00679  -0.1485   0.5667   0.4674
   1.250   0.7989   0.01487   0.00668  -0.1473   0.5204   0.4720
   1.500   0.8200   0.01514   0.00669  -0.1458   0.4757   0.4760
   1.750   0.8425   0.01551   0.00679  -0.1448   0.4469   0.4812
   2.000   0.8667   0.01582   0.00691  -0.1441   0.4279   0.4860
   2.250   0.8911   0.01609   0.00709  -0.1435   0.4145   0.4901
   2.500   0.9162   0.01640   0.00728  -0.1430   0.4039   0.4942
   2.750   0.9428   0.01660   0.00743  -0.1428   0.3947   0.4978
   3.000   0.9685   0.01683   0.00759  -0.1425   0.3877   0.5008
   3.250   0.9947   0.01699   0.00778  -0.1422   0.3808   0.5041
   3.500   1.0207   0.01727   0.00801  -0.1419   0.3749   0.5072
   3.750   1.0471   0.01744   0.00821  -0.1417   0.3684   0.5104
   4.000   1.0726   0.01769   0.00840  -0.1413   0.3609   0.5136
   4.250   1.0980   0.01783   0.00861  -0.1409   0.3538   0.5168
   4.500   1.1232   0.01802   0.00883  -0.1405   0.3472   0.5199
   4.750   1.1486   0.01824   0.00908  -0.1401   0.3410   0.5231
   5.000   1.1740   0.01842   0.00930  -0.1396   0.3349   0.5264
   5.250   1.1990   0.01868   0.00956  -0.1392   0.3300   0.5296
   5.500   1.2229   0.01875   0.00979  -0.1385   0.3216   0.5334
   5.750   1.2457   0.01890   0.00996  -0.1376   0.3108   0.5375
   6.000   1.2692   0.01907   0.01016  -0.1369   0.3017   0.5413
   6.250   1.2931   0.01918   0.01041  -0.1362   0.2940   0.5448
   6.500   1.3156   0.01934   0.01063  -0.1353   0.2831   0.5490
   6.750   1.3381   0.01946   0.01081  -0.1344   0.2669   0.5536
   7.000   1.3608   0.01965   0.01105  -0.1335   0.2518   0.5584
   7.250   1.3818   0.01994   0.01132  -0.1324   0.2280   0.5640
   7.500   1.3961   0.02084   0.01186  -0.1304   0.1655   0.5695
   7.750   1.4045   0.02242   0.01310  -0.1277   0.1155   0.5742
   8.000   1.4196   0.02334   0.01400  -0.1257   0.0941   0.5804
   8.250   1.4267   0.02483   0.01531  -0.1226   0.0555   0.5863
   8.500   1.4404   0.02559   0.01625  -0.1203   0.0483   0.5939
   8.750   1.4518   0.02645   0.01727  -0.1177   0.0411   0.6029
   9.000   1.4656   0.02720   0.01817  -0.1154   0.0281   0.6161
   9.250   1.4745   0.02827   0.01929  -0.1128   0.0237   0.6392
   9.500   1.4785   0.02889   0.02041  -0.1090   0.0221   1.0000
   9.750   1.4855   0.03020   0.02183  -0.1062   0.0210   1.0000
  10.000   1.4902   0.03169   0.02347  -0.1033   0.0203   1.0000
  10.250   1.4919   0.03347   0.02542  -0.1003   0.0198   1.0000
  10.500   1.4916   0.03548   0.02762  -0.0976   0.0195   1.0000
  10.750   1.4891   0.03782   0.03014  -0.0951   0.0194   1.0000
  11.000   1.4845   0.04054   0.03304  -0.0929   0.0192   1.0000
  11.250   1.4780   0.04370   0.03637  -0.0914   0.0190   1.0000
  11.500   1.4704   0.04728   0.04012  -0.0904   0.0188   1.0000
  11.750   1.4618   0.05126   0.04426  -0.0899   0.0186   1.0000
  12.000   1.4524   0.05555   0.04870  -0.0898   0.0183   1.0000
  12.250   1.4425   0.05999   0.05327  -0.0899   0.0181   1.0000
  12.500   1.4323   0.06454   0.05796  -0.0901   0.0178   1.0000
  12.750   1.4225   0.06910   0.06264  -0.0903   0.0176   1.0000
  13.000   1.4141   0.07348   0.06712  -0.0904   0.0175   1.0000
  13.250   1.4072   0.07766   0.07140  -0.0905   0.0175   1.0000
  13.500   1.4025   0.08148   0.07532  -0.0903   0.0174   1.0000
  13.750   1.4007   0.08477   0.07871  -0.0898   0.0174   1.0000
  14.000   1.4021   0.08749   0.08150  -0.0887   0.0174   1.0000
  14.250   1.4064   0.08975   0.08384  -0.0872   0.0175   1.0000
  14.500   1.4128   0.09171   0.08590  -0.0854   0.0176   1.0000
<< Back to GOE 165 (MVA MK.11) AIRFOIL (goe165-il)

Polar data table (+)

Polar graphs


<< Back to GOE 165 (MVA MK.11) AIRFOIL (goe165-il)