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

GOE 316 (HANSA-BRANDENBURG IV.5) AIRFOIL (goe316-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 316 (HANSA-BRANDENBURG IV.5) AIRFOIL (goe316-il)
Reynolds number: 200,000
Max Cl/Cd: 72.04 at α=9°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe316-il-200000-n5.txt
Download as CSV file: xf-goe316-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 316 (HANSA-BRANDENBURG IV.5) 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
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.4189   0.11694   0.11353   0.0001   1.0000   0.0245
  -9.500  -0.4163   0.11350   0.11011  -0.0030   1.0000   0.0246
  -9.250  -0.4126   0.10985   0.10650  -0.0056   1.0000   0.0247
  -9.000  -0.4080   0.10609   0.10276  -0.0079   1.0000   0.0247
  -8.500  -0.3957   0.09819   0.09493  -0.0103   1.0000   0.0249
  -8.250  -0.3850   0.09502   0.09176  -0.0092   1.0000   0.0254
  -8.000  -0.3780   0.09195   0.08873  -0.0104   1.0000   0.0259
  -7.750  -0.3718   0.08892   0.08573  -0.0120   1.0000   0.0267
  -7.500  -0.3656   0.08589   0.08275  -0.0146   1.0000   0.0285
  -7.250  -0.3521   0.08109   0.07795  -0.0267   1.0000   0.0305
  -7.000  -0.3300   0.07570   0.07247  -0.0352   0.9678   0.0307
  -6.750  -0.3057   0.07035   0.06699  -0.0418   0.9336   0.0308
  -6.500  -0.2851   0.06554   0.06201  -0.0461   0.8999   0.0308
  -6.250  -0.2761   0.06170   0.05813  -0.0456   0.8707   0.0314
  -6.000  -0.2648   0.05994   0.05632  -0.0439   0.8458   0.0327
  -5.750  -0.2468   0.05728   0.05351  -0.0453   0.8267   0.0349
  -5.500  -0.2103   0.05303   0.04874  -0.0517   0.8115   0.0385
  -5.250  -0.1866   0.04939   0.04477  -0.0534   0.7974   0.0386
  -4.750  -0.1496   0.04230   0.03745  -0.0549   0.7708   0.0401
  -4.500  -0.1275   0.04002   0.03502  -0.0554   0.7583   0.0412
  -4.250  -0.1028   0.03758   0.03234  -0.0561   0.7460   0.0423
  -4.000  -0.0763   0.03492   0.02938  -0.0567   0.7339   0.0425
  -3.500  -0.0172   0.02885   0.02237  -0.0572   0.7132   0.0379
  -3.250   0.0097   0.02676   0.01999  -0.0574   0.7028   0.0377
  -3.000   0.0392   0.02538   0.01813  -0.0571   0.6918   0.0389
  -2.750   0.0658   0.02320   0.01570  -0.0573   0.6804   0.0394
  -2.500   0.0940   0.02204   0.01422  -0.0571   0.6681   0.0390
  -2.250   0.1219   0.02080   0.01271  -0.0570   0.6548   0.0389
  -2.000   0.1498   0.01952   0.01118  -0.0570   0.6408   0.0391
  -1.750   0.1776   0.01840   0.00986  -0.0569   0.6261   0.0394
  -1.500   0.2053   0.01753   0.00882  -0.0569   0.6103   0.0400
  -1.250   0.2327   0.01698   0.00817  -0.0568   0.5934   0.0414
  -1.000   0.2605   0.01645   0.00749  -0.0567   0.5758   0.0425
  -0.750   0.2883   0.01586   0.00671  -0.0565   0.5579   0.0427
  -0.500   0.3160   0.01536   0.00605  -0.0563   0.5400   0.0430
  -0.250   0.3435   0.01495   0.00550  -0.0561   0.5217   0.0436
   0.000   0.3709   0.01462   0.00504  -0.0559   0.5030   0.0443
   0.250   0.3983   0.01435   0.00465  -0.0557   0.4840   0.0452
   0.500   0.4257   0.01413   0.00434  -0.0556   0.4644   0.0462
   0.750   0.4532   0.01407   0.00417  -0.0555   0.4452   0.0480
   1.000   0.4804   0.01382   0.00388  -0.0555   0.4279   0.0503
   1.250   0.5078   0.01371   0.00371  -0.0555   0.4127   0.0518
   1.500   0.5353   0.01367   0.00360  -0.0555   0.4002   0.0536
   1.750   0.5629   0.01369   0.00352  -0.0554   0.3903   0.0561
   2.000   0.5904   0.01375   0.00348  -0.0554   0.3820   0.0591
   2.250   0.6181   0.01377   0.00348  -0.0554   0.3755   0.0662
   2.500   0.6456   0.01379   0.00355  -0.0554   0.3698   0.0955
   3.000   0.6978   0.01218   0.00379  -0.0548   0.3607   1.0000
   3.250   0.7252   0.01239   0.00394  -0.0547   0.3563   1.0000
   3.500   0.7524   0.01263   0.00410  -0.0547   0.3525   1.0000
   3.750   0.7794   0.01290   0.00430  -0.0546   0.3494   1.0000
   4.000   0.8067   0.01314   0.00451  -0.0545   0.3467   1.0000
   4.250   0.8340   0.01336   0.00473  -0.0545   0.3440   1.0000
   4.500   0.8612   0.01361   0.00499  -0.0545   0.3413   1.0000
   4.750   0.8884   0.01387   0.00525  -0.0544   0.3389   1.0000
   5.000   0.9153   0.01416   0.00552  -0.0544   0.3366   1.0000
   5.250   0.9422   0.01447   0.00583  -0.0543   0.3343   1.0000
   5.500   0.9690   0.01475   0.00613  -0.0542   0.3304   1.0000
   5.750   0.9957   0.01497   0.00641  -0.0541   0.3244   1.0000
   6.000   1.0218   0.01527   0.00668  -0.0540   0.3187   1.0000
   6.250   1.0483   0.01556   0.00703  -0.0539   0.3146   1.0000
   6.500   1.0748   0.01580   0.00737  -0.0538   0.3103   1.0000
   6.750   1.1010   0.01606   0.00768  -0.0537   0.3053   1.0000
   7.000   1.1267   0.01639   0.00803  -0.0535   0.3002   1.0000
   7.250   1.1528   0.01659   0.00838  -0.0534   0.2940   1.0000
   7.500   1.1786   0.01687   0.00875  -0.0532   0.2896   1.0000
   7.750   1.2040   0.01722   0.00914  -0.0530   0.2858   1.0000
   8.000   1.2299   0.01745   0.00956  -0.0529   0.2804   1.0000
   8.250   1.2551   0.01770   0.00992  -0.0526   0.2734   1.0000
   8.500   1.2800   0.01791   0.01028  -0.0524   0.2630   1.0000
   8.750   1.3044   0.01811   0.01059  -0.0521   0.2459   1.0000
   9.000   1.3277   0.01843   0.01091  -0.0517   0.2164   1.0000
   9.250   1.3439   0.01948   0.01162  -0.0507   0.1629   1.0000
   9.500   1.3583   0.02086   0.01282  -0.0495   0.1281   1.0000
   9.750   1.3713   0.02230   0.01411  -0.0482   0.0934   1.0000
  10.000   1.3817   0.02387   0.01552  -0.0467   0.0668   1.0000
  10.250   1.3933   0.02520   0.01685  -0.0451   0.0572   1.0000
  10.500   1.4045   0.02644   0.01817  -0.0436   0.0497   1.0000
  10.750   1.4142   0.02770   0.01953  -0.0419   0.0427   1.0000
  11.000   1.4182   0.02912   0.02102  -0.0395   0.0362   1.0000
  11.250   1.4215   0.03075   0.02271  -0.0376   0.0300   1.0000
  11.500   1.4236   0.03267   0.02472  -0.0362   0.0256   1.0000
  11.750   1.4234   0.03504   0.02716  -0.0352   0.0229   1.0000
  12.000   1.4230   0.03767   0.02991  -0.0348   0.0210   1.0000
  12.250   1.4204   0.04073   0.03308  -0.0347   0.0196   1.0000
  12.500   1.4147   0.04434   0.03680  -0.0350   0.0186   1.0000
  12.750   1.4092   0.04808   0.04068  -0.0355   0.0180   1.0000
  13.000   1.4029   0.05206   0.04483  -0.0363   0.0175   1.0000
  13.250   1.3952   0.05636   0.04928  -0.0372   0.0170   1.0000
  13.500   1.3863   0.06090   0.05397  -0.0383   0.0166   1.0000
  13.750   1.3766   0.06566   0.05887  -0.0396   0.0163   1.0000
  14.000   1.3663   0.07051   0.06386  -0.0409   0.0159   1.0000
  14.250   1.3561   0.07542   0.06889  -0.0422   0.0156   1.0000
  14.500   1.3464   0.08041   0.07401  -0.0437   0.0153   1.0000
  14.750   1.3370   0.08542   0.07912  -0.0452   0.0150   1.0000
  15.000   1.3278   0.09046   0.08426  -0.0468   0.0147   1.0000
<< Back to GOE 316 (HANSA-BRANDENBURG IV.5) AIRFOIL (goe316-il)

Polar data table (+)

Polar graphs


<< Back to GOE 316 (HANSA-BRANDENBURG IV.5) AIRFOIL (goe316-il)