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

GOE 369 AIRFOIL (goe369-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 369 AIRFOIL (goe369-il)
Reynolds number: 200,000
Max Cl/Cd: 80.19 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe369-il-200000.txt
Download as CSV file: xf-goe369-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 369 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3434   0.09339   0.08999  -0.0113   1.0000   0.0190
  -7.500  -0.3398   0.09080   0.08744  -0.0119   1.0000   0.0195
  -7.250  -0.3371   0.08834   0.08503  -0.0126   1.0000   0.0202
  -7.000  -0.3304   0.08568   0.08242  -0.0145   1.0000   0.0210
  -6.750  -0.3192   0.08344   0.08021  -0.0187   1.0000   0.0218
  -6.500  -0.3035   0.08132   0.07810  -0.0246   1.0000   0.0221
  -6.250  -0.2860   0.07875   0.07551  -0.0290   1.0000   0.0222
  -6.000  -0.2683   0.07594   0.07268  -0.0323   1.0000   0.0223
  -5.500  -0.2516   0.06782   0.06465  -0.0321   1.0000   0.0227
  -5.250  -0.2461   0.06453   0.06144  -0.0307   1.0000   0.0231
  -5.000  -0.2411   0.06207   0.05905  -0.0298   1.0000   0.0235
  -4.750  -0.2027   0.05767   0.05459  -0.0364   0.9905   0.0245
  -4.500  -0.1575   0.05331   0.05013  -0.0442   0.9770   0.0258
  -4.250  -0.1102   0.04920   0.04587  -0.0517   0.9614   0.0275
  -4.000  -0.0464   0.04663   0.04291  -0.0602   0.9440   0.0295
  -3.750  -0.0047   0.04206   0.03810  -0.0653   0.9238   0.0301
  -3.500   0.0225   0.03766   0.03369  -0.0683   0.9014   0.0317
  -3.250   0.0538   0.03501   0.03084  -0.0702   0.8788   0.0338
  -3.000   0.0952   0.03503   0.03026  -0.0708   0.8585   0.0393
  -2.750   0.1140   0.03053   0.02568  -0.0714   0.8420   0.0413
  -2.500   0.1371   0.02860   0.02360  -0.0713   0.8271   0.0441
  -2.250   0.1713   0.02928   0.02368  -0.0703   0.8136   0.0511
  -2.000   0.1910   0.02525   0.01965  -0.0707   0.8014   0.0541
  -1.750   0.2218   0.02653   0.02038  -0.0694   0.7882   0.0640
  -1.500   0.2416   0.02248   0.01643  -0.0696   0.7760   0.0693
  -1.250   0.2676   0.02130   0.01498  -0.0690   0.7647   0.0815
  -1.000   0.2930   0.02014   0.01365  -0.0686   0.7544   0.0988
  -0.750   0.3173   0.01900   0.01236  -0.0684   0.7451   0.1353
   0.000   0.4085   0.01584   0.00810  -0.0647   0.7210   0.0621
   0.250   0.4358   0.01490   0.00703  -0.0640   0.7121   0.0612
   0.500   0.4628   0.01405   0.00605  -0.0632   0.7040   0.0582
   0.750   0.4892   0.01343   0.00536  -0.0624   0.6948   0.0581
   1.000   0.5151   0.01311   0.00501  -0.0616   0.6844   0.0632
   1.250   0.5403   0.01263   0.00452  -0.0608   0.6747   0.0661
   1.500   0.5661   0.01242   0.00429  -0.0600   0.6648   0.0748
   1.750   0.5919   0.01224   0.00408  -0.0593   0.6538   0.0875
   2.000   0.6358   0.01024   0.00388  -0.0627   0.6426   1.0000
   2.250   0.6615   0.01036   0.00382  -0.0619   0.6318   1.0000
   2.500   0.6870   0.01045   0.00379  -0.0612   0.6196   1.0000
   2.750   0.7124   0.01054   0.00381  -0.0605   0.6064   1.0000
   3.000   0.7378   0.01062   0.00383  -0.0597   0.5928   1.0000
   3.250   0.7630   0.01070   0.00386  -0.0590   0.5777   1.0000
   3.500   0.7881   0.01077   0.00386  -0.0582   0.5614   1.0000
   3.750   0.8132   0.01086   0.00391  -0.0574   0.5443   1.0000
   4.000   0.8382   0.01095   0.00398  -0.0567   0.5249   1.0000
   4.250   0.8631   0.01107   0.00408  -0.0559   0.5048   1.0000
   4.500   0.8878   0.01122   0.00419  -0.0552   0.4820   1.0000
   4.750   0.9121   0.01141   0.00432  -0.0544   0.4558   1.0000
   5.000   0.9358   0.01167   0.00448  -0.0536   0.4246   1.0000
   5.250   0.9586   0.01206   0.00470  -0.0526   0.3886   1.0000
   5.500   0.9805   0.01256   0.00506  -0.0516   0.3488   1.0000
   5.750   1.0018   0.01318   0.00549  -0.0506   0.3077   1.0000
   6.000   1.0228   0.01387   0.00598  -0.0496   0.2669   1.0000
   6.250   1.0436   0.01460   0.00652  -0.0486   0.2322   1.0000
   6.500   1.0645   0.01535   0.00715  -0.0476   0.2065   1.0000
   6.750   1.0845   0.01620   0.00784  -0.0466   0.1882   1.0000
   7.000   1.1046   0.01706   0.00861  -0.0455   0.1737   1.0000
   7.250   1.1259   0.01778   0.00934  -0.0446   0.1619   1.0000
   7.500   1.1472   0.01853   0.01013  -0.0437   0.1523   1.0000
   7.750   1.1675   0.01945   0.01101  -0.0427   0.1443   1.0000
   8.000   1.1894   0.02003   0.01170  -0.0419   0.1364   1.0000
   8.250   1.2098   0.02096   0.01261  -0.0410   0.1301   1.0000
   8.500   1.2314   0.02160   0.01340  -0.0401   0.1239   1.0000
   8.750   1.2514   0.02267   0.01442  -0.0392   0.1183   1.0000
   9.000   1.2724   0.02339   0.01536  -0.0382   0.1131   1.0000
   9.250   1.2926   0.02421   0.01624  -0.0373   0.1081   1.0000
   9.500   1.3127   0.02552   0.01759  -0.0364   0.1035   1.0000
   9.750   1.3314   0.02620   0.01851  -0.0353   0.0981   1.0000
  10.000   1.3495   0.02738   0.01959  -0.0344   0.0920   1.0000
  10.250   1.3650   0.02771   0.02023  -0.0329   0.0864   1.0000
  10.500   1.3804   0.02846   0.02101  -0.0317   0.0812   1.0000
  10.750   1.3947   0.02944   0.02219  -0.0303   0.0763   1.0000
  11.000   1.4076   0.03002   0.02295  -0.0287   0.0713   1.0000
  11.250   1.4178   0.03114   0.02412  -0.0270   0.0665   1.0000
  11.500   1.4248   0.03150   0.02476  -0.0245   0.0608   1.0000
  11.750   1.4280   0.03241   0.02575  -0.0222   0.0547   1.0000
  12.000   1.4298   0.03361   0.02703  -0.0203   0.0459   1.0000
  12.250   1.4283   0.03571   0.02923  -0.0187   0.0373   1.0000
  12.500   1.4239   0.03847   0.03214  -0.0174   0.0323   1.0000
  12.750   1.4184   0.04148   0.03521  -0.0169   0.0289   1.0000
  13.000   1.4111   0.04496   0.03882  -0.0165   0.0267   1.0000
  13.250   1.4088   0.04806   0.04212  -0.0165   0.0248   1.0000
  13.500   1.4043   0.05151   0.04573  -0.0170   0.0232   1.0000
  13.750   1.3973   0.05540   0.04975  -0.0176   0.0223   1.0000
  14.000   1.3872   0.05982   0.05427  -0.0186   0.0215   1.0000
  14.250   1.3748   0.06469   0.05925  -0.0196   0.0208   1.0000
  14.500   1.3647   0.06947   0.06419  -0.0208   0.0204   1.0000
  14.750   1.3553   0.07437   0.06930  -0.0224   0.0200   1.0000
  15.000   1.3446   0.07965   0.07477  -0.0244   0.0198   1.0000
  15.250   1.3329   0.08537   0.08069  -0.0268   0.0195   1.0000
  15.500   1.3197   0.09167   0.08718  -0.0297   0.0193   1.0000
  15.750   1.3060   0.09834   0.09404  -0.0331   0.0192   1.0000
  16.000   1.2910   0.10560   0.10148  -0.0369   0.0192   1.0000
  16.250   1.2746   0.11346   0.10953  -0.0413   0.0191   1.0000
  16.500   1.2570   0.12188   0.11812  -0.0461   0.0193   1.0000
  16.750   1.2379   0.13106   0.12747  -0.0516   0.0194   1.0000
  17.000   1.2176   0.14090   0.13746  -0.0576   0.0197   1.0000
  17.250   1.1953   0.15176   0.14843  -0.0643   0.0200   1.0000
  17.500   1.1711   0.16368   0.16045  -0.0716   0.0205   1.0000
  17.750   1.0531   0.21873   0.21524  -0.0996   0.0323   1.0000
  18.000   1.0561   0.22450   0.22099  -0.1024   0.0338   1.0000
  18.250   1.0618   0.22927   0.22577  -0.1044   0.0348   1.0000
  18.500   1.0740   0.22758   0.22414  -0.1042   0.0274   1.0000
  18.750   0.8548   0.22521   0.22233  -0.0923   0.0455   1.0000
  19.000   0.8568   0.22996   0.22710  -0.0942   0.0455   1.0000
<< Back to GOE 369 AIRFOIL (goe369-il)

Polar data table (+)

Polar graphs


<< Back to GOE 369 AIRFOIL (goe369-il)