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

GOE 532 AIRFOIL (goe532-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 532 AIRFOIL (goe532-il)
Reynolds number: 50,000
Max Cl/Cd: 34.68 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe532-il-50000-n5.txt
Download as CSV file: xf-goe532-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 532 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.2565   0.10042   0.09380  -0.0307   1.0000   0.1637
  -8.000  -0.2601   0.09842   0.09192  -0.0296   1.0000   0.1657
  -7.500  -0.3414   0.09229   0.08617  -0.0329   1.0000   0.1103
  -7.000  -0.3220   0.07961   0.07339  -0.0440   0.9853   0.0829
  -6.750  -0.2954   0.07458   0.06831  -0.0505   0.9728   0.0819
  -6.500  -0.2700   0.06859   0.06220  -0.0591   0.9590   0.0812
  -6.250  -0.2445   0.06264   0.05607  -0.0671   0.9448   0.0802
  -6.000  -0.2186   0.05679   0.04995  -0.0744   0.9305   0.0785
  -5.750  -0.1907   0.05097   0.04371  -0.0813   0.9165   0.0769
  -5.500  -0.1582   0.04583   0.03802  -0.0872   0.9043   0.0758
  -5.250  -0.1217   0.04216   0.03390  -0.0916   0.8934   0.0765
  -5.000  -0.0910   0.03947   0.03081  -0.0938   0.8792   0.0781
  -4.750  -0.0582   0.03694   0.02783  -0.0959   0.8660   0.0793
  -4.500  -0.0211   0.03455   0.02496  -0.0981   0.8551   0.0799
  -4.250   0.0098   0.03275   0.02277  -0.0988   0.8405   0.0806
  -4.000   0.0407   0.03126   0.02091  -0.0993   0.8263   0.0818
  -3.750   0.0730   0.02997   0.01927  -0.0997   0.8132   0.0840
  -3.500   0.1048   0.02877   0.01800  -0.1002   0.8009   0.0869
  -3.250   0.1323   0.02789   0.01699  -0.0997   0.7860   0.0894
  -3.000   0.1604   0.02709   0.01602  -0.0992   0.7722   0.0917
  -2.750   0.1900   0.02634   0.01507  -0.0989   0.7600   0.0946
  -2.500   0.2181   0.02567   0.01433  -0.0985   0.7475   0.0992
  -2.250   0.2442   0.02520   0.01378  -0.0980   0.7344   0.1054
  -2.000   0.2729   0.02470   0.01316  -0.0977   0.7234   0.1123
  -1.750   0.3016   0.02425   0.01262  -0.0976   0.7126   0.1233
  -1.500   0.3291   0.02377   0.01227  -0.0975   0.7015   0.1459
  -1.250   0.3590   0.02273   0.01192  -0.0980   0.6930   0.2842
  -1.000   0.3810   0.02230   0.01218  -0.0967   0.6823   0.4804
  -0.750   0.4012   0.02183   0.01227  -0.0939   0.6750   0.6485
  -0.500   0.4200   0.02131   0.01234  -0.0906   0.6657   0.8293
  -0.250   0.4664   0.02122   0.01196  -0.0937   0.6581   1.0000
   0.000   0.4928   0.02162   0.01209  -0.0938   0.6491   1.0000
   0.250   0.5221   0.02191   0.01208  -0.0940   0.6419   1.0000
   0.500   0.5491   0.02231   0.01225  -0.0939   0.6345   1.0000
   0.750   0.5758   0.02274   0.01249  -0.0938   0.6272   1.0000
   1.000   0.6053   0.02305   0.01257  -0.0939   0.6217   1.0000
   1.250   0.6291   0.02362   0.01305  -0.0935   0.6141   1.0000
   1.500   0.6558   0.02404   0.01332  -0.0933   0.6073   1.0000
   1.750   0.6839   0.02439   0.01349  -0.0931   0.6010   1.0000
   2.000   0.7062   0.02499   0.01405  -0.0924   0.5924   1.0000
   2.250   0.7351   0.02523   0.01411  -0.0922   0.5854   1.0000
   2.500   0.7565   0.02581   0.01466  -0.0913   0.5758   1.0000
   2.750   0.7845   0.02602   0.01473  -0.0908   0.5676   1.0000
   3.000   0.8059   0.02655   0.01524  -0.0899   0.5578   1.0000
   3.250   0.8339   0.02669   0.01525  -0.0893   0.5490   1.0000
   3.500   0.8542   0.02720   0.01575  -0.0882   0.5379   1.0000
   3.750   0.8833   0.02723   0.01562  -0.0877   0.5293   1.0000
   4.000   0.9017   0.02784   0.01631  -0.0864   0.5183   1.0000
   4.250   0.9299   0.02801   0.01636  -0.0859   0.5111   1.0000
   4.500   0.9479   0.02875   0.01720  -0.0847   0.5016   1.0000
   4.750   0.9758   0.02891   0.01730  -0.0843   0.4944   1.0000
   5.000   0.9929   0.02967   0.01816  -0.0829   0.4844   1.0000
   5.250   1.0205   0.02981   0.01823  -0.0824   0.4767   1.0000
   5.500   1.0364   0.03060   0.01915  -0.0809   0.4663   1.0000
   5.750   1.0644   0.03069   0.01919  -0.0803   0.4586   1.0000
   6.000   1.0784   0.03157   0.02022  -0.0787   0.4474   1.0000
   6.250   1.1019   0.03190   0.02055  -0.0777   0.4385   1.0000
   6.500   1.1192   0.03255   0.02130  -0.0763   0.4275   1.0000
   6.750   1.1364   0.03321   0.02204  -0.0748   0.4168   1.0000
   7.000   1.1599   0.03345   0.02228  -0.0738   0.4070   1.0000
   7.250   1.1710   0.03441   0.02338  -0.0718   0.3942   1.0000
   7.500   1.1857   0.03510   0.02413  -0.0701   0.3821   1.0000
   7.750   1.2045   0.03544   0.02446  -0.0686   0.3707   1.0000
   8.000   1.2144   0.03634   0.02543  -0.0664   0.3579   1.0000
   8.250   1.2219   0.03739   0.02657  -0.0642   0.3455   1.0000
   8.500   1.2319   0.03820   0.02739  -0.0620   0.3347   1.0000
   8.750   1.2432   0.03893   0.02811  -0.0600   0.3247   1.0000
   9.000   1.2454   0.04046   0.02973  -0.0576   0.3149   1.0000
   9.250   1.2608   0.04111   0.03036  -0.0561   0.3073   1.0000
   9.500   1.2610   0.04303   0.03240  -0.0541   0.2988   1.0000
   9.750   1.2753   0.04381   0.03314  -0.0527   0.2916   1.0000
  10.000   1.2739   0.04601   0.03546  -0.0511   0.2836   1.0000
  10.250   1.2871   0.04688   0.03628  -0.0498   0.2764   1.0000
  10.500   1.2837   0.04947   0.03903  -0.0486   0.2690   1.0000
  10.750   1.3014   0.05005   0.03953  -0.0475   0.2624   1.0000
  11.000   1.2925   0.05347   0.04316  -0.0467   0.2559   1.0000
  11.250   1.3069   0.05431   0.04395  -0.0457   0.2486   1.0000
  11.500   1.2970   0.05801   0.04781  -0.0453   0.2416   1.0000
  11.750   1.3081   0.05929   0.04908  -0.0445   0.2344   1.0000
  12.000   1.3021   0.06287   0.05283  -0.0444   0.2283   1.0000
  12.250   1.3023   0.06568   0.05574  -0.0442   0.2216   1.0000
  12.500   1.3074   0.06791   0.05801  -0.0438   0.2150   1.0000
  12.750   1.2937   0.07279   0.06308  -0.0446   0.2081   1.0000
  13.000   1.3117   0.07323   0.06344  -0.0435   0.2010   1.0000
  13.250   1.2807   0.08090   0.07142  -0.0458   0.1943   1.0000
  13.500   1.3076   0.07993   0.07035  -0.0440   0.1872   1.0000
  13.750   1.2631   0.09026   0.08099  -0.0478   0.1808   1.0000
  14.000   1.2522   0.09552   0.08637  -0.0493   0.1742   1.0000
  14.250   1.2684   0.09619   0.08702  -0.0484   0.1686   1.0000
  14.500   1.1581   0.12113   0.11229  -0.0610   0.1587   1.0000
  14.750   1.1893   0.11850   0.10967  -0.0584   0.1557   1.0000
<< Back to GOE 532 AIRFOIL (goe532-il)

Polar data table (+)

Polar graphs


<< Back to GOE 532 AIRFOIL (goe532-il)