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

GOE 567 AIRFOIL (goe567-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 567 AIRFOIL (goe567-il)
Reynolds number: 200,000
Max Cl/Cd: 79.79 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe567-il-200000-n5.txt
Download as CSV file: xf-goe567-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 567 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
  ------ -------- --------- --------- -------- -------- --------
 -12.250  -0.6609   0.05024   0.04614  -0.0951   0.9918   0.0463
 -12.000  -0.6825   0.03979   0.03525  -0.1111   0.9738   0.0472
 -11.750  -0.6643   0.03553   0.03074  -0.1201   0.9619   0.0485
 -11.500  -0.6313   0.03464   0.02982  -0.1238   0.9571   0.0499
 -11.250  -0.6077   0.03329   0.02835  -0.1265   0.9467   0.0514
 -11.000  -0.5831   0.03134   0.02610  -0.1301   0.9377   0.0535
 -10.750  -0.5586   0.02912   0.02345  -0.1335   0.9283   0.0556
 -10.500  -0.5303   0.02864   0.02300  -0.1345   0.9189   0.0571
 -10.250  -0.4973   0.02812   0.02245  -0.1366   0.9114   0.0588
 -10.000  -0.4707   0.02722   0.02137  -0.1378   0.9002   0.0607
  -9.750  -0.4427   0.02595   0.01981  -0.1395   0.8902   0.0626
  -9.500  -0.4147   0.02480   0.01832  -0.1408   0.8795   0.0642
  -9.250  -0.3884   0.02372   0.01711  -0.1416   0.8681   0.0657
  -9.000  -0.3603   0.02294   0.01624  -0.1423   0.8578   0.0669
  -8.750  -0.3329   0.02223   0.01541  -0.1428   0.8471   0.0682
  -8.500  -0.3067   0.02159   0.01462  -0.1430   0.8367   0.0698
  -8.250  -0.2796   0.02095   0.01379  -0.1433   0.8276   0.0716
  -8.000  -0.2540   0.02035   0.01299  -0.1433   0.8177   0.0733
  -7.750  -0.2273   0.01978   0.01221  -0.1434   0.8096   0.0746
  -7.500  -0.2018   0.01915   0.01144  -0.1433   0.8009   0.0759
  -7.250  -0.1758   0.01848   0.01072  -0.1433   0.7937   0.0777
  -7.000  -0.1502   0.01802   0.01021  -0.1431   0.7857   0.0795
  -6.750  -0.1235   0.01761   0.00969  -0.1431   0.7789   0.0814
  -6.500  -0.0973   0.01720   0.00920  -0.1429   0.7719   0.0832
  -6.250  -0.0708   0.01682   0.00871  -0.1428   0.7652   0.0850
  -6.000  -0.0437   0.01651   0.00825  -0.1427   0.7592   0.0869
  -5.750  -0.0176   0.01611   0.00781  -0.1426   0.7525   0.0889
  -5.500   0.0089   0.01570   0.00737  -0.1425   0.7464   0.0913
  -5.250   0.0357   0.01539   0.00701  -0.1424   0.7407   0.0936
  -5.000   0.0624   0.01512   0.00669  -0.1423   0.7343   0.0960
  -4.750   0.0898   0.01490   0.00638  -0.1422   0.7287   0.0989
  -4.500   0.1172   0.01470   0.00610  -0.1421   0.7232   0.1016
  -4.250   0.1439   0.01440   0.00581  -0.1420   0.7169   0.1053
  -4.000   0.1714   0.01419   0.00555  -0.1420   0.7114   0.1092
  -3.750   0.1989   0.01402   0.00533  -0.1420   0.7060   0.1139
  -3.500   0.2261   0.01381   0.00514  -0.1419   0.6998   0.1207
  -3.250   0.2538   0.01364   0.00493  -0.1419   0.6943   0.1309
  -3.000   0.2812   0.01343   0.00477  -0.1419   0.6888   0.1479
  -2.750   0.3084   0.01322   0.00465  -0.1418   0.6826   0.1753
  -2.500   0.3359   0.01304   0.00453  -0.1419   0.6772   0.2088
  -2.250   0.3633   0.01290   0.00447  -0.1418   0.6715   0.2428
  -2.000   0.3905   0.01280   0.00446  -0.1417   0.6652   0.2737
  -1.750   0.4181   0.01275   0.00443  -0.1416   0.6598   0.2999
  -1.500   0.4456   0.01273   0.00444  -0.1415   0.6540   0.3222
  -1.250   0.4728   0.01272   0.00445  -0.1413   0.6477   0.3416
  -1.000   0.5004   0.01273   0.00444  -0.1411   0.6422   0.3617
  -0.750   0.5275   0.01275   0.00451  -0.1409   0.6361   0.3830
  -0.500   0.5547   0.01278   0.00456  -0.1407   0.6300   0.4015
   0.000   0.6088   0.01286   0.00463  -0.1402   0.6168   0.4317
   0.250   0.6357   0.01291   0.00467  -0.1399   0.6102   0.4463
   0.500   0.6623   0.01295   0.00474  -0.1395   0.6031   0.4609
   0.750   0.6889   0.01301   0.00479  -0.1392   0.5959   0.4742
   1.000   0.7155   0.01307   0.00484  -0.1388   0.5888   0.4871
   1.250   0.7419   0.01313   0.00492  -0.1385   0.5818   0.4984
   1.500   0.7686   0.01321   0.00497  -0.1382   0.5761   0.5089
   1.750   0.7949   0.01329   0.00508  -0.1378   0.5694   0.5204
   2.000   0.8210   0.01336   0.00516  -0.1374   0.5625   0.5313
   2.250   0.8470   0.01345   0.00526  -0.1370   0.5557   0.5423
   2.500   0.8729   0.01354   0.00537  -0.1366   0.5489   0.5541
   2.750   0.8990   0.01364   0.00548  -0.1362   0.5436   0.5657
   3.000   0.9248   0.01374   0.00564  -0.1358   0.5376   0.5777
   3.250   0.9505   0.01384   0.00578  -0.1353   0.5316   0.5905
   3.750   1.0014   0.01406   0.00611  -0.1344   0.5204   0.6196
   4.000   1.0258   0.01417   0.00626  -0.1337   0.5130   0.6367
   4.250   1.0493   0.01426   0.00643  -0.1328   0.5049   0.6569
   4.500   1.0725   0.01435   0.00659  -0.1319   0.4960   0.6839
   4.750   1.0949   0.01437   0.00677  -0.1308   0.4875   0.7306
   5.000   1.1181   0.01418   0.00689  -0.1296   0.4795   1.0000
   5.250   1.1416   0.01441   0.00711  -0.1289   0.4710   1.0000
   5.500   1.1643   0.01466   0.00732  -0.1280   0.4622   1.0000
   5.750   1.1872   0.01490   0.00757  -0.1271   0.4534   1.0000
   6.000   1.2090   0.01517   0.00780  -0.1261   0.4450   1.0000
   6.250   1.2311   0.01543   0.00809  -0.1251   0.4363   1.0000
   6.500   1.2516   0.01573   0.00837  -0.1239   0.4276   1.0000
   6.750   1.2724   0.01602   0.00868  -0.1227   0.4178   1.0000
   7.000   1.2916   0.01634   0.00901  -0.1212   0.4082   1.0000
   7.250   1.3088   0.01669   0.00935  -0.1194   0.3969   1.0000
   7.500   1.3254   0.01704   0.00971  -0.1175   0.3859   1.0000
   7.750   1.3403   0.01745   0.01011  -0.1153   0.3745   1.0000
   8.000   1.3529   0.01793   0.01058  -0.1129   0.3603   1.0000
   8.250   1.3649   0.01848   0.01110  -0.1104   0.3458   1.0000
   8.500   1.3754   0.01911   0.01170  -0.1078   0.3295   1.0000
   8.750   1.3846   0.01983   0.01239  -0.1051   0.3123   1.0000
   9.000   1.3896   0.02078   0.01325  -0.1020   0.2908   1.0000
   9.250   1.3916   0.02195   0.01433  -0.0988   0.2681   1.0000
   9.500   1.3941   0.02321   0.01552  -0.0958   0.2470   1.0000
   9.750   1.3961   0.02461   0.01686  -0.0930   0.2278   1.0000
  10.000   1.3961   0.02625   0.01842  -0.0903   0.2080   1.0000
  10.250   1.3961   0.02802   0.02013  -0.0879   0.1869   1.0000
  10.500   1.3945   0.03002   0.02205  -0.0856   0.1652   1.0000
  10.750   1.3920   0.03222   0.02418  -0.0834   0.1474   1.0000
  11.000   1.3904   0.03446   0.02638  -0.0816   0.1357   1.0000
  11.250   1.3925   0.03651   0.02844  -0.0801   0.1286   1.0000
  11.500   1.3930   0.03875   0.03070  -0.0787   0.1226   1.0000
  11.750   1.3971   0.04076   0.03277  -0.0776   0.1172   1.0000
  12.000   1.3992   0.04300   0.03506  -0.0765   0.1122   1.0000
  12.250   1.3993   0.04552   0.03761  -0.0755   0.1077   1.0000
  12.500   1.4060   0.04745   0.03964  -0.0748   0.1024   1.0000
  12.750   1.4086   0.04985   0.04210  -0.0741   0.0970   1.0000
  13.000   1.4123   0.05219   0.04451  -0.0735   0.0921   1.0000
  13.250   1.4188   0.05429   0.04671  -0.0730   0.0851   1.0000
  13.500   1.4221   0.05679   0.04925  -0.0726   0.0755   1.0000
  13.750   1.4220   0.05974   0.05218  -0.0723   0.0696   1.0000
  14.000   1.4216   0.06281   0.05523  -0.0721   0.0655   1.0000
  14.250   1.4217   0.06590   0.05835  -0.0720   0.0629   1.0000
  14.500   1.4221   0.06901   0.06154  -0.0720   0.0608   1.0000
  14.750   1.4223   0.07221   0.06480  -0.0720   0.0592   1.0000
  15.000   1.4212   0.07563   0.06828  -0.0722   0.0574   1.0000
  15.250   1.4193   0.07922   0.07192  -0.0725   0.0556   1.0000
  15.500   1.4217   0.08227   0.07509  -0.0728   0.0538   1.0000
  15.750   1.4228   0.08552   0.07845  -0.0732   0.0518   1.0000
  16.000   1.4225   0.08904   0.08205  -0.0737   0.0502   1.0000
  16.250   1.4202   0.09286   0.08594  -0.0744   0.0486   1.0000
<< Back to GOE 567 AIRFOIL (goe567-il)

Polar data table (+)

Polar graphs


<< Back to GOE 567 AIRFOIL (goe567-il)