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

GOE 116 (MVA MK.3) AIRFOIL (goe116-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 116 (MVA MK.3) AIRFOIL (goe116-il)
Reynolds number: 200,000
Max Cl/Cd: 78.44 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe116-il-200000.txt
Download as CSV file: xf-goe116-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 116 (MVA MK.3) 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3772   0.09338   0.09026  -0.0348   1.0000   0.0687
  -8.250  -0.3094   0.07950   0.07683  -0.0471   0.9948   0.0779
  -8.000  -0.3668   0.08732   0.08429  -0.0368   1.0000   0.0731
  -7.000  -0.3515   0.03644   0.03260  -0.0949   0.9620   0.0436
  -6.750  -0.3238   0.02848   0.02373  -0.0996   0.9516   0.0352
  -6.500  -0.2890   0.02596   0.02042  -0.1011   0.9416   0.0307
  -6.250  -0.2629   0.02237   0.01635  -0.1018   0.9294   0.0292
  -6.000  -0.2358   0.02021   0.01384  -0.1017   0.9175   0.0281
  -5.750  -0.2093   0.01863   0.01201  -0.1013   0.9063   0.0274
  -5.500  -0.1831   0.01741   0.01057  -0.1007   0.8960   0.0274
  -5.250  -0.1571   0.01649   0.00948  -0.1002   0.8864   0.0279
  -5.000  -0.1305   0.01581   0.00859  -0.0998   0.8759   0.0291
  -4.750  -0.1035   0.01517   0.00770  -0.0995   0.8669   0.0325
  -4.500  -0.0762   0.01451   0.00681  -0.0991   0.8591   0.0364
  -4.250  -0.0489   0.01372   0.00596  -0.0989   0.8505   0.0529
  -4.000  -0.0204   0.01422   0.00652  -0.0987   0.8433   0.1412
  -3.750   0.0077   0.01516   0.00741  -0.0986   0.8350   0.1600
  -3.500   0.0350   0.01574   0.00790  -0.0985   0.8284   0.1743
  -3.250   0.0627   0.01593   0.00788  -0.0984   0.8210   0.1818
  -3.000   0.0895   0.01601   0.00796  -0.0984   0.8151   0.1927
  -2.750   0.1167   0.01602   0.00796  -0.0985   0.8081   0.2031
  -2.500   0.1441   0.01600   0.00776  -0.0984   0.8029   0.2112
  -2.250   0.1720   0.01566   0.00738  -0.0986   0.7961   0.2132
  -2.000   0.1997   0.01543   0.00705  -0.0986   0.7907   0.2156
  -1.750   0.2276   0.01522   0.00677  -0.0987   0.7850   0.2172
  -1.500   0.2556   0.01502   0.00650  -0.0988   0.7792   0.2187
  -1.250   0.2834   0.01487   0.00625  -0.0987   0.7747   0.2203
  -1.000   0.3115   0.01474   0.00609  -0.0989   0.7685   0.2220
  -0.750   0.3394   0.01464   0.00591  -0.0989   0.7635   0.2235
  -0.500   0.3670   0.01443   0.00573  -0.0990   0.7586   0.2259
  -0.250   0.3947   0.01430   0.00567  -0.0991   0.7529   0.2290
   0.000   0.4225   0.01421   0.00558  -0.0991   0.7485   0.2321
   0.250   0.4503   0.01417   0.00557  -0.0992   0.7433   0.2350
   0.500   0.4782   0.01414   0.00555  -0.0993   0.7381   0.2378
   0.750   0.5058   0.01401   0.00549  -0.0993   0.7341   0.2415
   1.000   0.5334   0.01400   0.00560  -0.0995   0.7286   0.2461
   1.250   0.5612   0.01399   0.00564  -0.0995   0.7236   0.2514
   1.500   0.5890   0.01392   0.00565  -0.0996   0.7199   0.2585
   1.750   0.6166   0.01395   0.00583  -0.0998   0.7142   0.2682
   2.000   0.6443   0.01390   0.00595  -0.0999   0.7093   0.2860
   2.250   0.6687   0.01223   0.00607  -0.0990   0.7058   1.0000
   2.500   0.6961   0.01244   0.00627  -0.0990   0.6993   1.0000
   2.750   0.7240   0.01253   0.00631  -0.0988   0.6937   1.0000
   3.000   0.7511   0.01268   0.00651  -0.0987   0.6867   1.0000
   3.250   0.7787   0.01277   0.00661  -0.0985   0.6804   1.0000
   3.500   0.8053   0.01270   0.00657  -0.0979   0.6692   1.0000
   3.750   0.8310   0.01224   0.00598  -0.0966   0.6494   1.0000
   4.000   0.8564   0.01203   0.00581  -0.0957   0.6305   1.0000
   4.250   0.8814   0.01176   0.00556  -0.0945   0.6046   1.0000
   4.500   0.9029   0.01151   0.00514  -0.0925   0.5390   1.0000
   4.750   0.9037   0.01380   0.00580  -0.0882   0.2468   1.0000
   5.000   0.9054   0.01673   0.00758  -0.0849   0.0442   1.0000
   5.250   0.9260   0.01756   0.00849  -0.0836   0.0369   1.0000
   5.500   0.9443   0.01859   0.00965  -0.0821   0.0320   1.0000
   5.750   0.9623   0.01957   0.01081  -0.0805   0.0301   1.0000
   6.000   0.9775   0.02076   0.01210  -0.0785   0.0289   1.0000
   6.250   0.9908   0.02207   0.01349  -0.0762   0.0282   1.0000
   6.500   1.0041   0.02349   0.01496  -0.0739   0.0279   1.0000
   6.750   1.0196   0.02503   0.01654  -0.0717   0.0280   1.0000
   7.000   1.0396   0.02662   0.01815  -0.0702   0.0274   1.0000
   7.250   1.0614   0.02823   0.01977  -0.0693   0.0253   1.0000
   7.500   1.0934   0.03046   0.02215  -0.0690   0.0266   1.0000
   7.750   1.1329   0.03446   0.02640  -0.0694   0.0312   1.0000
  16.000   0.7270   0.17689   0.17391  -0.0641   0.0427   1.0000
  16.250   0.7215   0.18107   0.17808  -0.0670   0.0405   1.0000
<< Back to GOE 116 (MVA MK.3) AIRFOIL (goe116-il)

Polar data table (+)

Polar graphs


<< Back to GOE 116 (MVA MK.3) AIRFOIL (goe116-il)