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

GOE 621 AIRFOIL (goe621-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 621 AIRFOIL (goe621-il)
Reynolds number: 200,000
Max Cl/Cd: 70.06 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe621-il-200000-n5.txt
Download as CSV file: xf-goe621-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 621 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
  ------ -------- --------- --------- -------- -------- --------
 -11.750  -0.4530   0.05872   0.05430  -0.1029   0.9725   0.0393
 -11.500  -0.4949   0.04379   0.03884  -0.1250   0.9602   0.0391
 -11.250  -0.5094   0.03913   0.03377  -0.1290   0.9481   0.0393
 -11.000  -0.5007   0.03628   0.03066  -0.1308   0.9415   0.0395
 -10.750  -0.4936   0.03460   0.02885  -0.1299   0.9318   0.0398
 -10.500  -0.4761   0.03289   0.02699  -0.1304   0.9260   0.0401
 -10.250  -0.4669   0.03156   0.02553  -0.1288   0.9162   0.0404
 -10.000  -0.4490   0.03017   0.02397  -0.1286   0.9101   0.0408
  -9.750  -0.4366   0.02900   0.02265  -0.1270   0.9008   0.0411
  -9.500  -0.4181   0.02778   0.02126  -0.1264   0.8940   0.0416
  -9.250  -0.4019   0.02675   0.02008  -0.1250   0.8852   0.0422
  -9.000  -0.3826   0.02568   0.01881  -0.1242   0.8777   0.0429
  -8.750  -0.3634   0.02469   0.01763  -0.1232   0.8700   0.0436
  -8.500  -0.3432   0.02376   0.01650  -0.1222   0.8620   0.0442
  -8.250  -0.3205   0.02285   0.01539  -0.1215   0.8554   0.0448
  -8.000  -0.2999   0.02201   0.01449  -0.1205   0.8464   0.0453
  -7.750  -0.2753   0.02121   0.01364  -0.1201   0.8401   0.0459
  -7.500  -0.2539   0.02059   0.01298  -0.1190   0.8307   0.0465
  -7.250  -0.2293   0.01994   0.01223  -0.1185   0.8240   0.0474
  -7.000  -0.2063   0.01937   0.01159  -0.1177   0.8159   0.0482
  -6.750  -0.1818   0.01881   0.01092  -0.1171   0.8087   0.0492
  -6.500  -0.1571   0.01832   0.01030  -0.1165   0.8016   0.0505
  -6.250  -0.1333   0.01778   0.00971  -0.1157   0.7938   0.0517
  -6.000  -0.1079   0.01729   0.00916  -0.1153   0.7877   0.0531
  -5.750  -0.0842   0.01689   0.00873  -0.1145   0.7802   0.0545
  -5.500  -0.0588   0.01649   0.00825  -0.1140   0.7741   0.0562
  -5.250  -0.0331   0.01614   0.00780  -0.1135   0.7685   0.0579
  -5.000  -0.0091   0.01575   0.00741  -0.1128   0.7617   0.0602
  -4.750   0.0169   0.01545   0.00703  -0.1124   0.7560   0.0631
  -4.500   0.0428   0.01515   0.00670  -0.1119   0.7507   0.0670
  -4.250   0.0679   0.01490   0.00645  -0.1113   0.7450   0.0722
  -4.000   0.0940   0.01465   0.00620  -0.1109   0.7398   0.0792
  -3.500   0.1464   0.01427   0.00581  -0.1101   0.7297   0.1011
  -3.250   0.1724   0.01410   0.00564  -0.1096   0.7245   0.1140
  -3.000   0.1995   0.01393   0.00545  -0.1094   0.7194   0.1272
  -2.750   0.2251   0.01378   0.00532  -0.1088   0.7131   0.1411
  -2.500   0.2505   0.01363   0.00520  -0.1083   0.7062   0.1572
  -2.250   0.2774   0.01348   0.00506  -0.1080   0.7004   0.1784
  -2.000   0.3025   0.01336   0.00501  -0.1073   0.6942   0.2033
  -1.750   0.3279   0.01323   0.00498  -0.1068   0.6876   0.2332
  -1.500   0.3546   0.01313   0.00493  -0.1064   0.6821   0.2674
  -1.250   0.3795   0.01307   0.00494  -0.1057   0.6753   0.2951
  -1.000   0.4049   0.01302   0.00492  -0.1051   0.6682   0.3183
  -0.750   0.4316   0.01299   0.00487  -0.1047   0.6620   0.3383
  -0.500   0.4562   0.01297   0.00489  -0.1039   0.6549   0.3558
  -0.250   0.4823   0.01296   0.00487  -0.1034   0.6490   0.3725
   0.000   0.5088   0.01295   0.00486  -0.1029   0.6436   0.3892
   0.250   0.5334   0.01295   0.00490  -0.1022   0.6371   0.4063
   0.500   0.5590   0.01293   0.00491  -0.1016   0.6312   0.4249
   0.750   0.5847   0.01291   0.00492  -0.1010   0.6254   0.4460
   1.000   0.6082   0.01287   0.00499  -0.1000   0.6180   0.4719
   1.250   0.6328   0.01281   0.00499  -0.0992   0.6111   0.5031
   1.500   0.6559   0.01273   0.00506  -0.0982   0.6039   0.5400
   1.750   0.6780   0.01258   0.00510  -0.0969   0.5963   0.5933
   2.000   0.7612   0.01198   0.00532  -0.1079   0.5852   0.9692
   2.250   0.8019   0.01213   0.00536  -0.1106   0.5747   1.0000
   2.500   0.8229   0.01225   0.00544  -0.1092   0.5642   1.0000
   2.750   0.8439   0.01239   0.00550  -0.1077   0.5535   1.0000
   3.000   0.8639   0.01253   0.00556  -0.1061   0.5406   1.0000
   3.250   0.8831   0.01270   0.00565  -0.1043   0.5255   1.0000
   3.500   0.9015   0.01288   0.00575  -0.1023   0.5088   1.0000
   3.750   0.9185   0.01311   0.00586  -0.1001   0.4901   1.0000
   4.000   0.9340   0.01338   0.00600  -0.0977   0.4690   1.0000
   4.250   0.9480   0.01370   0.00618  -0.0950   0.4469   1.0000
   4.500   0.9601   0.01408   0.00640  -0.0920   0.4253   1.0000
   4.750   0.9718   0.01448   0.00666  -0.0890   0.4069   1.0000
   5.000   0.9818   0.01489   0.00695  -0.0857   0.3906   1.0000
   5.250   0.9927   0.01530   0.00726  -0.0826   0.3770   1.0000
   5.500   1.0044   0.01574   0.00761  -0.0798   0.3652   1.0000
   5.750   1.0166   0.01621   0.00798  -0.0771   0.3541   1.0000
   6.000   1.0306   0.01665   0.00837  -0.0748   0.3438   1.0000
   6.250   1.0429   0.01718   0.00880  -0.0723   0.3344   1.0000
   6.500   1.0589   0.01761   0.00920  -0.0705   0.3257   1.0000
   6.750   1.0729   0.01813   0.00966  -0.0684   0.3184   1.0000
   7.000   1.0896   0.01857   0.01009  -0.0668   0.3121   1.0000
   7.250   1.1055   0.01905   0.01056  -0.0650   0.3055   1.0000
   7.500   1.1201   0.01961   0.01107  -0.0632   0.2999   1.0000
   7.750   1.1376   0.02005   0.01154  -0.0618   0.2946   1.0000
   8.000   1.1539   0.02056   0.01206  -0.0603   0.2891   1.0000
   8.250   1.1689   0.02114   0.01261  -0.0586   0.2841   1.0000
   8.500   1.1850   0.02169   0.01318  -0.0571   0.2793   1.0000
   8.750   1.2013   0.02222   0.01375  -0.0557   0.2738   1.0000
   9.000   1.2158   0.02285   0.01438  -0.0542   0.2684   1.0000
   9.250   1.2297   0.02353   0.01505  -0.0526   0.2633   1.0000
   9.500   1.2456   0.02412   0.01570  -0.0513   0.2577   1.0000
   9.750   1.2590   0.02485   0.01644  -0.0497   0.2521   1.0000
  10.000   1.2717   0.02564   0.01722  -0.0482   0.2473   1.0000
  10.250   1.2870   0.02632   0.01798  -0.0470   0.2419   1.0000
  10.500   1.3000   0.02713   0.01881  -0.0456   0.2365   1.0000
  10.750   1.3106   0.02809   0.01977  -0.0440   0.2317   1.0000
  11.000   1.3251   0.02887   0.02063  -0.0429   0.2263   1.0000
  11.250   1.3367   0.02983   0.02162  -0.0415   0.2208   1.0000
  11.500   1.3459   0.03096   0.02274  -0.0401   0.2160   1.0000
  11.750   1.3591   0.03188   0.02375  -0.0390   0.2107   1.0000
  12.000   1.3690   0.03302   0.02493  -0.0377   0.2052   1.0000
  12.250   1.3770   0.03433   0.02623  -0.0364   0.2004   1.0000
  12.500   1.3885   0.03544   0.02742  -0.0354   0.1950   1.0000
  12.750   1.3970   0.03678   0.02880  -0.0342   0.1900   1.0000
  13.000   1.4037   0.03829   0.03030  -0.0331   0.1859   1.0000
  13.250   1.4145   0.03954   0.03165  -0.0322   0.1815   1.0000
  13.500   1.4229   0.04099   0.03315  -0.0312   0.1771   1.0000
  13.750   1.4285   0.04269   0.03486  -0.0302   0.1730   1.0000
  14.000   1.4361   0.04426   0.03649  -0.0294   0.1691   1.0000
  14.250   1.4439   0.04587   0.03818  -0.0286   0.1649   1.0000
  14.500   1.4488   0.04775   0.04009  -0.0279   0.1608   1.0000
  14.750   1.4524   0.04977   0.04211  -0.0271   0.1572   1.0000
  15.000   1.4601   0.05148   0.04395  -0.0266   0.1535   1.0000
  15.250   1.4650   0.05349   0.04602  -0.0260   0.1496   1.0000
  15.500   1.4682   0.05568   0.04825  -0.0255   0.1464   1.0000
  15.750   1.4708   0.05796   0.05054  -0.0251   0.1438   1.0000
  16.000   1.4773   0.05992   0.05263  -0.0248   0.1411   1.0000
  16.250   1.4816   0.06214   0.05496  -0.0245   0.1382   1.0000
  16.500   1.4845   0.06453   0.05744  -0.0243   0.1354   1.0000
  16.750   1.4850   0.06722   0.06018  -0.0242   0.1328   1.0000
  17.000   1.4858   0.06992   0.06291  -0.0242   0.1302   1.0000
  17.250   1.4887   0.07248   0.06562  -0.0243   0.1276   1.0000
  17.500   1.4901   0.07524   0.06851  -0.0245   0.1252   1.0000
  17.750   1.4894   0.07830   0.07166  -0.0248   0.1225   1.0000
  18.000   1.4866   0.08168   0.07510  -0.0253   0.1199   1.0000
  18.250   1.4841   0.08505   0.07852  -0.0258   0.1175   1.0000
  18.500   1.4830   0.08836   0.08200  -0.0265   0.1151   1.0000
  18.750   1.4801   0.09197   0.08575  -0.0273   0.1124   1.0000
  19.000   1.4749   0.09594   0.08982  -0.0284   0.1097   1.0000
  19.250   1.4676   0.10025   0.09418  -0.0297   0.1070   1.0000
<< Back to GOE 621 AIRFOIL (goe621-il)

Polar data table (+)

Polar graphs


<< Back to GOE 621 AIRFOIL (goe621-il)