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

GOE 619 AIRFOIL (goe619-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 619 AIRFOIL (goe619-il)
Reynolds number: 100,000
Max Cl/Cd: 52.4 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe619-il-100000.txt
Download as CSV file: xf-goe619-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 619 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3377   0.10548   0.10099  -0.0234   1.0000   0.1500
  -8.000  -0.3684   0.10517   0.10079  -0.0217   1.0000   0.1532
  -7.750  -0.4225   0.10469   0.10046  -0.0270   0.9944   0.1549
  -7.500  -0.3687   0.09919   0.09491  -0.0244   0.9943   0.1578
  -7.250  -0.3421   0.09591   0.09159  -0.0269   0.9891   0.1623
  -6.750  -0.4432   0.04861   0.04287  -0.0750   0.9577   0.0971
  -6.500  -0.4285   0.04337   0.03698  -0.0775   0.9500   0.0975
  -6.250  -0.4023   0.04044   0.03395  -0.0794   0.9441   0.0998
  -6.000  -0.3709   0.03923   0.03272  -0.0811   0.9385   0.1033
  -5.750  -0.3493   0.03649   0.02949  -0.0818   0.9308   0.1060
  -5.500  -0.3147   0.03366   0.02586  -0.0843   0.9263   0.1103
  -5.250  -0.2944   0.03224   0.02440  -0.0837   0.9187   0.1143
  -5.000  -0.2617   0.03121   0.02311  -0.0850   0.9128   0.1234
  -4.750  -0.2210   0.03004   0.02188  -0.0876   0.9091   0.1357
  -4.500  -0.2063   0.02945   0.02125  -0.0857   0.8998   0.1455
  -4.250  -0.1690   0.02899   0.02074  -0.0875   0.8948   0.1616
  -4.000  -0.1448   0.02882   0.02048  -0.0871   0.8874   0.1740
  -3.750  -0.1148   0.02866   0.02013  -0.0876   0.8807   0.1865
  -3.500  -0.0732   0.02861   0.02010  -0.0900   0.8766   0.2030
  -3.250  -0.0593   0.02891   0.02056  -0.0879   0.8670   0.2126
  -3.000  -0.0218   0.02857   0.02017  -0.0896   0.8619   0.2254
  -2.750   0.0101   0.02823   0.01964  -0.0903   0.8558   0.2367
  -2.500   0.0336   0.02806   0.01955  -0.0896   0.8476   0.2454
  -2.250   0.0769   0.02746   0.01892  -0.0919   0.8436   0.2608
  -2.000   0.0938   0.02752   0.01895  -0.0902   0.8341   0.2744
  -1.750   0.1317   0.02708   0.01857  -0.0916   0.8286   0.2969
  -1.500   0.1817   0.02635   0.01789  -0.0948   0.8256   0.3282
  -1.000   0.2411   0.02597   0.01756  -0.0948   0.8101   0.3775
  -0.750   0.2594   0.02608   0.01771  -0.0930   0.7996   0.3956
  -0.500   0.3083   0.02530   0.01700  -0.0957   0.7946   0.4219
  -0.250   0.3642   0.02419   0.01593  -0.0993   0.7912   0.4515
   0.000   0.3804   0.02420   0.01600  -0.0970   0.7787   0.4707
   0.250   0.4344   0.02290   0.01480  -0.1001   0.7745   0.5107
   0.500   0.4525   0.02272   0.01480  -0.0980   0.7624   0.5404
   0.750   0.5035   0.02137   0.01371  -0.1007   0.7576   0.5980
   1.000   0.5200   0.02090   0.01386  -0.0980   0.7458   0.7140
   1.250   0.6198   0.01944   0.01251  -0.1101   0.7406   1.0000
   1.500   0.6386   0.01963   0.01258  -0.1081   0.7282   1.0000
   1.750   0.6696   0.01947   0.01226  -0.1080   0.7185   1.0000
   2.000   0.7024   0.01921   0.01186  -0.1080   0.7083   1.0000
   2.250   0.7249   0.01925   0.01183  -0.1065   0.6957   1.0000
   2.500   0.7548   0.01909   0.01155  -0.1062   0.6845   1.0000
   2.750   0.7876   0.01885   0.01117  -0.1062   0.6733   1.0000
   3.000   0.8104   0.01886   0.01113  -0.1047   0.6589   1.0000
   3.250   0.8350   0.01882   0.01101  -0.1035   0.6441   1.0000
   3.500   0.8603   0.01875   0.01086  -0.1024   0.6285   1.0000
   3.750   0.8853   0.01870   0.01072  -0.1012   0.6119   1.0000
   4.000   0.9096   0.01869   0.01061  -0.0999   0.5942   1.0000
   4.250   0.9336   0.01871   0.01051  -0.0986   0.5755   1.0000
   4.500   0.9578   0.01876   0.01043  -0.0974   0.5560   1.0000
   4.750   0.9776   0.01895   0.01052  -0.0955   0.5344   1.0000
   5.000   0.9972   0.01919   0.01064  -0.0936   0.5126   1.0000
   5.250   1.0185   0.01948   0.01077  -0.0921   0.4922   1.0000
   5.500   1.0406   0.01986   0.01097  -0.0908   0.4736   1.0000
   5.750   1.0626   0.02035   0.01130  -0.0896   0.4567   1.0000
   6.000   1.0842   0.02090   0.01176  -0.0884   0.4414   1.0000
   6.250   1.1058   0.02151   0.01229  -0.0873   0.4277   1.0000
   6.500   1.1282   0.02214   0.01285  -0.0863   0.4156   1.0000
   6.750   1.1537   0.02277   0.01332  -0.0859   0.4049   1.0000
   7.000   1.1739   0.02340   0.01397  -0.0847   0.3942   1.0000
   7.250   1.1963   0.02408   0.01462  -0.0838   0.3849   1.0000
   7.500   1.2199   0.02469   0.01515  -0.0831   0.3757   1.0000
   7.750   1.2390   0.02538   0.01591  -0.0818   0.3671   1.0000
   8.000   1.2634   0.02600   0.01645  -0.0813   0.3589   1.0000
   8.250   1.2822   0.02675   0.01728  -0.0799   0.3518   1.0000
   8.500   1.3036   0.02744   0.01801  -0.0790   0.3450   1.0000
   8.750   1.3291   0.02820   0.01871  -0.0788   0.3386   1.0000
   9.000   1.3428   0.02892   0.01962  -0.0766   0.3319   1.0000
   9.250   1.3695   0.02956   0.02016  -0.0766   0.3253   1.0000
   9.500   1.3840   0.03040   0.02118  -0.0747   0.3196   1.0000
   9.750   1.3999   0.03112   0.02200  -0.0729   0.3134   1.0000
  10.000   1.4280   0.03179   0.02255  -0.0732   0.3068   1.0000
  10.250   1.4324   0.03257   0.02360  -0.0697   0.3006   1.0000
  10.500   1.4529   0.03312   0.02415  -0.0687   0.2939   1.0000
  10.750   1.4666   0.03392   0.02504  -0.0668   0.2878   1.0000
  11.000   1.4736   0.03469   0.02598  -0.0639   0.2818   1.0000
  11.250   1.5003   0.03514   0.02628  -0.0638   0.2744   1.0000
  11.500   1.4924   0.03599   0.02743  -0.0587   0.2687   1.0000
  11.750   1.5043   0.03638   0.02778  -0.0564   0.2613   1.0000
  12.000   1.5059   0.03721   0.02874  -0.0530   0.2551   1.0000
  12.250   1.5036   0.03799   0.02964  -0.0491   0.2484   1.0000
  12.500   1.5142   0.03855   0.03016  -0.0471   0.2412   1.0000
  12.750   1.5042   0.03984   0.03170  -0.0430   0.2351   1.0000
  13.000   1.5181   0.04021   0.03193  -0.0415   0.2272   1.0000
  13.250   1.5009   0.04208   0.03412  -0.0374   0.2212   1.0000
  13.500   1.5025   0.04312   0.03518  -0.0351   0.2139   1.0000
  13.750   1.4934   0.04501   0.03725  -0.0325   0.2070   1.0000
  14.000   1.4889   0.04671   0.03903  -0.0304   0.1995   1.0000
  14.250   1.4798   0.04897   0.04142  -0.0285   0.1921   1.0000
  14.500   1.4725   0.05121   0.04374  -0.0270   0.1843   1.0000
  14.750   1.4615   0.05411   0.04678  -0.0258   0.1767   1.0000
  15.000   1.4546   0.05673   0.04946  -0.0249   0.1693   1.0000
  15.250   1.4427   0.06025   0.05312  -0.0243   0.1621   1.0000
  15.500   1.4361   0.06331   0.05622  -0.0241   0.1557   1.0000
  15.750   1.4267   0.06702   0.06003  -0.0241   0.1498   1.0000
  16.000   1.4181   0.07079   0.06390  -0.0244   0.1444   1.0000
  16.250   1.4169   0.07364   0.06670  -0.0245   0.1400   1.0000
  16.500   1.4034   0.07853   0.07180  -0.0254   0.1357   1.0000
  16.750   1.3963   0.08258   0.07595  -0.0262   0.1319   1.0000
  17.000   1.3982   0.08524   0.07857  -0.0266   0.1283   1.0000
  17.250   1.3865   0.09027   0.08377  -0.0280   0.1255   1.0000
  17.500   1.3719   0.09592   0.08962  -0.0298   0.1228   1.0000
  17.750   1.3617   0.10095   0.09478  -0.0316   0.1201   1.0000
  18.000   1.3642   0.10376   0.09758  -0.0325   0.1169   1.0000
  18.250   1.3582   0.10816   0.10206  -0.0341   0.1142   1.0000
  18.500   1.3328   0.11629   0.11043  -0.0378   0.1127   1.0000
  18.750   1.2908   0.12798   0.12238  -0.0437   0.1121   1.0000
<< Back to GOE 619 AIRFOIL (goe619-il)

Polar data table (+)

Polar graphs


<< Back to GOE 619 AIRFOIL (goe619-il)