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

GOE 417 AIRFOIL (goe417-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 417 AIRFOIL (goe417-il)
Reynolds number: 100,000
Max Cl/Cd: 62.02 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe417-il-100000.txt
Download as CSV file: xf-goe417-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 417 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  -0.2893   0.10072   0.09616  -0.0211   1.0000   0.0410
  -7.250  -0.2983   0.09977   0.09530  -0.0189   1.0000   0.0416
  -7.000  -0.3066   0.09879   0.09441  -0.0171   1.0000   0.0420
  -6.750  -0.3121   0.09809   0.09379  -0.0171   1.0000   0.0427
  -6.500  -0.3144   0.09742   0.09317  -0.0185   1.0000   0.0431
  -6.250  -0.3108   0.09709   0.09288  -0.0224   1.0000   0.0435
  -6.000  -0.3005   0.09617   0.09195  -0.0270   1.0000   0.0437
  -5.750  -0.3085   0.08979   0.08569  -0.0201   1.0000   0.0446
  -5.500  -0.3097   0.08618   0.08211  -0.0161   1.0000   0.0459
  -5.250  -0.3069   0.08361   0.07958  -0.0152   1.0000   0.0475
  -5.000  -0.3010   0.08119   0.07718  -0.0156   1.0000   0.0492
  -4.750  -0.2922   0.07873   0.07473  -0.0171   1.0000   0.0510
  -4.500  -0.2782   0.07632   0.07231  -0.0199   1.0000   0.0534
  -4.250  -0.2252   0.07636   0.07206  -0.0347   1.0000   0.0562
  -4.000  -0.2258   0.07119   0.06701  -0.0324   1.0000   0.0570
  -3.750  -0.2260   0.06768   0.06359  -0.0293   1.0000   0.0584
  -3.500  -0.2102   0.06473   0.06064  -0.0303   0.9986   0.0612
  -3.250  -0.1283   0.06158   0.05709  -0.0475   0.9909   0.0707
  -3.000  -0.1099   0.05691   0.05253  -0.0479   0.9846   0.0753
  -2.750  -0.0449   0.05363   0.04894  -0.0594   0.9758   0.0857
  -2.500  -0.0034   0.04997   0.04524  -0.0643   0.9669   0.0953
  -2.250   0.0514   0.04641   0.04148  -0.0719   0.9571   0.1083
  -2.000   0.1059   0.04366   0.03844  -0.0791   0.9460   0.1296
  -1.750   0.1485   0.04068   0.03539  -0.0834   0.9379   0.1466
  -1.500   0.1949   0.03858   0.03306  -0.0882   0.9292   0.1739
   0.000   0.4692   0.02642   0.01935  -0.1078   0.8773   0.1883
   0.250   0.5257   0.02378   0.01596  -0.1104   0.8713   0.1182
   0.500   0.5650   0.02239   0.01427  -0.1112   0.8603   0.1075
   0.750   0.6078   0.02100   0.01258  -0.1125   0.8511   0.1014
   1.000   0.6484   0.01977   0.01132  -0.1138   0.8407   0.1051
   1.250   0.6832   0.01880   0.01031  -0.1139   0.8268   0.1068
   1.500   0.7170   0.01789   0.00936  -0.1137   0.8119   0.1084
   1.750   0.7488   0.01721   0.00860  -0.1132   0.7940   0.1121
   2.000   0.7782   0.01664   0.00801  -0.1124   0.7728   0.1205
   2.250   0.8094   0.01609   0.00746  -0.1118   0.7514   0.1481
   2.500   0.8429   0.01442   0.00707  -0.1120   0.7252   1.0000
   2.750   0.8696   0.01443   0.00678  -0.1108   0.6964   1.0000
   3.000   0.8958   0.01457   0.00664  -0.1097   0.6666   1.0000
   3.250   0.9211   0.01486   0.00663  -0.1086   0.6366   1.0000
   3.500   0.9452   0.01524   0.00673  -0.1073   0.6065   1.0000
   3.750   0.9683   0.01566   0.00692  -0.1060   0.5777   1.0000
   4.000   0.9910   0.01608   0.00720  -0.1048   0.5512   1.0000
   4.250   1.0144   0.01652   0.00748  -0.1037   0.5288   1.0000
   4.500   1.0379   0.01697   0.00781  -0.1028   0.5089   1.0000
   4.750   1.0618   0.01743   0.00817  -0.1019   0.4914   1.0000
   5.000   1.0858   0.01792   0.00855  -0.1011   0.4753   1.0000
   5.250   1.1097   0.01843   0.00902  -0.1003   0.4598   1.0000
   5.500   1.1335   0.01896   0.00950  -0.0995   0.4451   1.0000
   5.750   1.1571   0.01952   0.01004  -0.0988   0.4307   1.0000
   6.000   1.1806   0.02010   0.01061  -0.0980   0.4164   1.0000
   6.250   1.2038   0.02071   0.01127  -0.0971   0.4022   1.0000
   6.500   1.2269   0.02136   0.01194  -0.0963   0.3879   1.0000
   6.750   1.2497   0.02206   0.01266  -0.0954   0.3734   1.0000
   7.000   1.2723   0.02280   0.01343  -0.0945   0.3589   1.0000
   7.250   1.2945   0.02358   0.01431  -0.0936   0.3448   1.0000
   7.500   1.3165   0.02441   0.01522  -0.0927   0.3313   1.0000
   7.750   1.3377   0.02518   0.01608  -0.0916   0.3178   1.0000
   8.000   1.3568   0.02567   0.01664  -0.0902   0.3028   1.0000
   8.250   1.3740   0.02594   0.01699  -0.0884   0.2869   1.0000
   8.500   1.3899   0.02617   0.01733  -0.0865   0.2718   1.0000
   8.750   1.4038   0.02632   0.01759  -0.0843   0.2559   1.0000
   9.000   1.4169   0.02652   0.01796  -0.0820   0.2402   1.0000
   9.250   1.4273   0.02666   0.01838  -0.0792   0.2190   1.0000
   9.500   1.4330   0.02702   0.01878  -0.0759   0.1805   1.0000
   9.750   1.4299   0.02899   0.02028  -0.0717   0.1052   1.0000
  10.000   1.4231   0.03151   0.02232  -0.0674   0.0672   1.0000
  10.250   1.4159   0.03394   0.02468  -0.0632   0.0596   1.0000
  10.500   1.4126   0.03609   0.02699  -0.0598   0.0548   1.0000
  10.750   1.4064   0.03852   0.02958  -0.0565   0.0515   1.0000
  11.000   1.3975   0.04131   0.03247  -0.0536   0.0495   1.0000
  11.250   1.3884   0.04437   0.03562  -0.0510   0.0480   1.0000
  11.500   1.3891   0.04682   0.03822  -0.0491   0.0467   1.0000
  11.750   1.3914   0.04933   0.04088  -0.0474   0.0454   1.0000
  12.000   1.3968   0.05182   0.04349  -0.0459   0.0441   1.0000
  12.250   1.4036   0.05436   0.04615  -0.0445   0.0423   1.0000
  12.500   1.4111   0.05702   0.04891  -0.0432   0.0403   1.0000
  13.000   1.4667   0.06463   0.05665  -0.0415   0.0373   1.0000
  13.250   1.4680   0.06839   0.06071  -0.0404   0.0372   1.0000
  13.500   1.4653   0.07254   0.06517  -0.0394   0.0372   1.0000
  13.750   1.4602   0.07718   0.07008  -0.0387   0.0373   1.0000
  14.000   1.4488   0.08114   0.07435  -0.0383   0.0375   1.0000
  14.250   1.4347   0.08536   0.07883  -0.0385   0.0376   1.0000
  14.500   1.4186   0.08993   0.08367  -0.0393   0.0378   1.0000
  14.750   1.4004   0.09501   0.08902  -0.0411   0.0380   1.0000
  15.000   1.3799   0.10075   0.09503  -0.0438   0.0383   1.0000
  15.250   1.3572   0.10735   0.10190  -0.0476   0.0387   1.0000
  15.500   1.3321   0.11512   0.10994  -0.0527   0.0391   1.0000
  15.750   1.3053   0.12423   0.11931  -0.0591   0.0396   1.0000
  16.000   1.2703   0.13622   0.13155  -0.0682   0.0404   1.0000
  16.250   1.2182   0.15490   0.15046  -0.0827   0.0420   1.0000
<< Back to GOE 417 AIRFOIL (goe417-il)

Polar data table (+)

Polar graphs


<< Back to GOE 417 AIRFOIL (goe417-il)