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

GOE 336 (MVA H.44) AIRFOIL (goe336-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 336 (MVA H.44) AIRFOIL (goe336-il)
Reynolds number: 200,000
Max Cl/Cd: 79.53 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe336-il-200000.txt
Download as CSV file: xf-goe336-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 336 (MVA H.44) 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3111   0.10503   0.10168  -0.0309   1.0000   0.0383
  -8.500  -0.3173   0.10326   0.09998  -0.0333   1.0000   0.0385
  -8.250  -0.3214   0.10110   0.09789  -0.0350   1.0000   0.0386
  -8.000  -0.2974   0.09378   0.09055  -0.0300   1.0000   0.0397
  -7.750  -0.2949   0.09123   0.08805  -0.0290   1.0000   0.0405
  -7.500  -0.2991   0.08927   0.08617  -0.0276   1.0000   0.0411
  -7.250  -0.3059   0.08752   0.08450  -0.0260   1.0000   0.0418
  -7.000  -0.3119   0.08577   0.08285  -0.0249   0.9990   0.0426
  -6.750  -0.2796   0.08088   0.07790  -0.0341   0.9906   0.0454
  -6.500  -0.2326   0.07457   0.07141  -0.0538   0.9761   0.0484
  -6.250  -0.2115   0.07017   0.06705  -0.0535   0.9709   0.0497
  -6.000  -0.1842   0.06653   0.06338  -0.0570   0.9586   0.0517
  -5.750  -0.1528   0.06272   0.05948  -0.0627   0.9435   0.0552
  -5.500  -0.1084   0.05797   0.05437  -0.0742   0.9241   0.0594
  -5.250  -0.0939   0.05455   0.05098  -0.0731   0.9056   0.0607
  -5.000  -0.0751   0.05210   0.04846  -0.0730   0.8854   0.0629
  -4.750  -0.0528   0.04960   0.04579  -0.0740   0.8655   0.0664
  -4.500  -0.0176   0.04659   0.04226  -0.0784   0.8461   0.0717
  -4.250  -0.0030   0.04358   0.03923  -0.0774   0.8283   0.0728
  -4.000   0.0152   0.04147   0.03702  -0.0767   0.8116   0.0746
  -3.750   0.0367   0.03966   0.03503  -0.0765   0.7958   0.0785
  -3.500   0.0685   0.03742   0.03224  -0.0778   0.7816   0.0857
  -3.250   0.0874   0.03503   0.02979  -0.0772   0.7675   0.0871
  -3.000   0.1089   0.03331   0.02792  -0.0767   0.7540   0.0894
  -2.750   0.1422   0.03376   0.02770  -0.0763   0.7410   0.0990
  -2.500   0.1613   0.02994   0.02384  -0.0762   0.7289   0.1007
  -2.250   0.1834   0.02824   0.02210  -0.0757   0.7164   0.1031
  -2.000   0.2083   0.02716   0.02084  -0.0752   0.7047   0.1094
  -1.750   0.2399   0.02185   0.01444  -0.0737   0.6967   0.0729
  -1.500   0.2663   0.02016   0.01249  -0.0730   0.6853   0.0677
  -1.250   0.2931   0.01859   0.01056  -0.0723   0.6752   0.0656
  -1.000   0.3204   0.01741   0.00904  -0.0717   0.6659   0.0651
  -0.750   0.3476   0.01674   0.00819  -0.0713   0.6552   0.0671
  -0.500   0.3751   0.01615   0.00741  -0.0709   0.6463   0.0689
  -0.250   0.4025   0.01558   0.00668  -0.0705   0.6375   0.0701
   0.000   0.4294   0.01497   0.00601  -0.0701   0.6295   0.0724
   0.250   0.4561   0.01463   0.00568  -0.0697   0.6220   0.0769
   0.500   0.4829   0.01443   0.00542  -0.0693   0.6148   0.0841
   0.750   0.5090   0.01410   0.00518  -0.0688   0.6080   0.1025
   1.000   0.5350   0.01391   0.00503  -0.0684   0.6021   0.1416
   1.250   0.5611   0.01381   0.00502  -0.0681   0.5952   0.1644
   1.500   0.5877   0.01379   0.00498  -0.0678   0.5900   0.1872
   1.750   0.6132   0.01369   0.00506  -0.0674   0.5837   0.2269
   2.000   0.7066   0.01223   0.00510  -0.0817   0.5755   1.0000
   2.250   0.7322   0.01241   0.00521  -0.0812   0.5697   1.0000
   2.500   0.7578   0.01258   0.00531  -0.0807   0.5641   1.0000
   2.750   0.7837   0.01278   0.00537  -0.0803   0.5592   1.0000
   3.000   0.8086   0.01294   0.00555  -0.0797   0.5526   1.0000
   3.250   0.8339   0.01308   0.00561  -0.0791   0.5459   1.0000
   3.500   0.8586   0.01323   0.00573  -0.0785   0.5383   1.0000
   3.750   0.8836   0.01337   0.00580  -0.0778   0.5311   1.0000
   4.000   0.9083   0.01355   0.00598  -0.0772   0.5244   1.0000
   4.250   0.9329   0.01368   0.00609  -0.0765   0.5169   1.0000
   4.500   0.9575   0.01384   0.00623  -0.0759   0.5096   1.0000
   4.750   0.9817   0.01396   0.00634  -0.0751   0.5013   1.0000
   5.000   1.0058   0.01411   0.00651  -0.0744   0.4934   1.0000
   5.250   1.0300   0.01424   0.00663  -0.0736   0.4854   1.0000
   5.500   1.0538   0.01440   0.00683  -0.0729   0.4775   1.0000
   5.750   1.0779   0.01455   0.00700  -0.0722   0.4703   1.0000
   6.000   1.1015   0.01471   0.00723  -0.0714   0.4625   1.0000
   6.250   1.1250   0.01484   0.00737  -0.0705   0.4536   1.0000
   6.500   1.1473   0.01495   0.00756  -0.0695   0.4425   1.0000
   6.750   1.1693   0.01505   0.00772  -0.0684   0.4296   1.0000
   7.000   1.1905   0.01515   0.00786  -0.0671   0.4138   1.0000
   7.250   1.2112   0.01529   0.00801  -0.0658   0.3953   1.0000
   7.500   1.2318   0.01550   0.00823  -0.0645   0.3771   1.0000
   7.750   1.2518   0.01574   0.00852  -0.0632   0.3509   1.0000
   8.000   1.2676   0.01621   0.00883  -0.0612   0.3040   1.0000
   8.250   1.2767   0.01720   0.00947  -0.0584   0.2536   1.0000
   8.500   1.2871   0.01821   0.01029  -0.0558   0.2241   1.0000
   8.750   1.2982   0.01916   0.01112  -0.0534   0.1999   1.0000
   9.000   1.3114   0.01995   0.01187  -0.0514   0.1786   1.0000
   9.250   1.3228   0.02081   0.01264  -0.0491   0.1546   1.0000
   9.500   1.3307   0.02184   0.01352  -0.0464   0.1329   1.0000
   9.750   1.3375   0.02279   0.01441  -0.0434   0.1188   1.0000
  10.000   1.3454   0.02366   0.01529  -0.0407   0.1087   1.0000
  10.250   1.3531   0.02462   0.01626  -0.0381   0.0967   1.0000
  10.500   1.3595   0.02574   0.01738  -0.0357   0.0801   1.0000
  10.750   1.3618   0.02723   0.01878  -0.0331   0.0673   1.0000
  11.000   1.3642   0.02886   0.02037  -0.0309   0.0601   1.0000
  11.250   1.3695   0.03039   0.02197  -0.0292   0.0545   1.0000
  11.500   1.3718   0.03230   0.02393  -0.0276   0.0494   1.0000
  11.750   1.3789   0.03391   0.02566  -0.0265   0.0427   1.0000
  12.000   1.3837   0.03581   0.02765  -0.0254   0.0347   1.0000
  12.250   1.3829   0.03836   0.03017  -0.0245   0.0275   1.0000
  12.500   1.3825   0.04095   0.03280  -0.0237   0.0251   1.0000
  12.750   1.3821   0.04363   0.03556  -0.0230   0.0241   1.0000
  13.000   1.3811   0.04643   0.03848  -0.0225   0.0234   1.0000
  13.250   1.3790   0.04941   0.04159  -0.0221   0.0228   1.0000
  13.500   1.3758   0.05260   0.04491  -0.0218   0.0224   1.0000
  13.750   1.3715   0.05596   0.04841  -0.0217   0.0220   1.0000
  14.000   1.3660   0.05956   0.05214  -0.0218   0.0217   1.0000
  14.250   1.3594   0.06341   0.05612  -0.0220   0.0214   1.0000
  14.500   1.3518   0.06749   0.06033  -0.0224   0.0211   1.0000
  14.750   1.3434   0.07177   0.06473  -0.0229   0.0208   1.0000
  15.000   1.3346   0.07619   0.06927  -0.0236   0.0205   1.0000
  15.250   1.3282   0.08031   0.07352  -0.0242   0.0202   1.0000
  15.500   1.3230   0.08434   0.07767  -0.0248   0.0200   1.0000
  15.750   1.3182   0.08831   0.08175  -0.0254   0.0198   1.0000
  16.000   1.3145   0.09212   0.08567  -0.0260   0.0197   1.0000
<< Back to GOE 336 (MVA H.44) AIRFOIL (goe336-il)

Polar data table (+)

Polar graphs


<< Back to GOE 336 (MVA H.44) AIRFOIL (goe336-il)