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

GOE 393 AIRFOIL (goe393-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 393 AIRFOIL (goe393-il)
Reynolds number: 50,000
Max Cl/Cd: 34.68 at α=8.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe393-il-50000.txt
Download as CSV file: xf-goe393-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 393 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4628   0.11516   0.10842   0.0147   1.0000   0.1241
  -8.250  -0.4637   0.11394   0.10728   0.0114   1.0000   0.1275
  -8.000  -0.4686   0.11433   0.10779   0.0045   1.0000   0.1286
  -7.750  -0.4459   0.10558   0.09900   0.0102   1.0000   0.1335
  -7.500  -0.4391   0.10253   0.09599   0.0087   1.0000   0.1381
  -7.250  -0.4358   0.10131   0.09483   0.0022   1.0000   0.1422
  -7.000  -0.4256   0.09698   0.09056   0.0005   1.0000   0.1445
  -6.750  -0.4128   0.09249   0.08609   0.0019   1.0000   0.1494
  -6.500  -0.4013   0.09043   0.08405  -0.0046   1.0000   0.1563
  -6.250  -0.3896   0.08594   0.07961  -0.0049   1.0000   0.1606
  -6.000  -0.3745   0.08340   0.07706  -0.0093   1.0000   0.1708
  -5.750  -0.3619   0.07901   0.07271  -0.0091   1.0000   0.1763
  -5.500  -0.3418   0.07689   0.07053  -0.0165   1.0000   0.1872
  -5.250  -0.3304   0.07224   0.06598  -0.0134   1.0000   0.1937
  -5.000  -0.3121   0.06905   0.06278  -0.0170   1.0000   0.2046
  -4.750  -0.2931   0.06601   0.05972  -0.0199   1.0000   0.2181
  -4.500  -0.2769   0.06270   0.05645  -0.0206   1.0000   0.2338
  -4.250  -0.2617   0.05952   0.05332  -0.0204   1.0000   0.2528
  -4.000  -0.2441   0.05670   0.05052  -0.0221   1.0000   0.2779
  -3.750  -0.2285   0.05398   0.04785  -0.0225   1.0000   0.3063
  -3.500  -0.2178   0.05105   0.04505  -0.0209   1.0000   0.3369
  -3.250  -0.2120   0.04836   0.04254  -0.0179   1.0000   0.3710
  -3.000  -0.2140   0.04618   0.04055  -0.0140   1.0000   0.4139
  -2.750  -0.2205   0.04452   0.03905  -0.0098   1.0000   0.4477
  -2.500  -0.2240   0.04293   0.03754  -0.0069   1.0000   0.4855
  -2.250  -0.2251   0.04090   0.03561  -0.0032   1.0000   0.5188
  -2.000  -0.2211   0.03895   0.03374  -0.0006   1.0000   0.5557
  -1.750  -0.1875   0.03600   0.03080  -0.0029   0.9895   0.6071
  -1.500  -0.1111   0.03282   0.02740  -0.0162   0.9726   0.6198
  -1.250   0.1049   0.03287   0.02453  -0.0623   0.9502   0.2299
  -1.000   0.1667   0.03044   0.02161  -0.0686   0.9344   0.2100
  -0.750   0.2254   0.02871   0.01918  -0.0736   0.9183   0.2014
  -0.500   0.2727   0.02720   0.01741  -0.0770   0.9014   0.2165
  -0.250   0.3145   0.02583   0.01587  -0.0790   0.8847   0.2346
   0.000   0.3542   0.02487   0.01473  -0.0803   0.8681   0.2691
   0.250   0.3904   0.02417   0.01391  -0.0808   0.8518   0.2980
   0.500   0.4223   0.02377   0.01341  -0.0806   0.8357   0.3180
   0.750   0.4500   0.02350   0.01321  -0.0799   0.8195   0.3412
   1.000   0.4763   0.02334   0.01318  -0.0792   0.8040   0.3714
   1.250   0.5012   0.02298   0.01324  -0.0783   0.7895   0.4334
   1.500   0.5354   0.02235   0.01334  -0.0781   0.7748   1.0000
   1.750   0.5612   0.02308   0.01372  -0.0773   0.7603   1.0000
   2.000   0.5859   0.02387   0.01428  -0.0765   0.7462   1.0000
   2.250   0.6101   0.02473   0.01499  -0.0760   0.7326   1.0000
   2.500   0.6341   0.02568   0.01584  -0.0756   0.7199   1.0000
   2.750   0.6581   0.02664   0.01673  -0.0752   0.7080   1.0000
   3.000   0.6826   0.02752   0.01757  -0.0744   0.6974   1.0000
   3.250   0.7062   0.02863   0.01868  -0.0743   0.6863   1.0000
   3.500   0.7289   0.02996   0.02005  -0.0745   0.6753   1.0000
   3.750   0.7523   0.03112   0.02123  -0.0742   0.6657   1.0000
   4.000   0.7754   0.03238   0.02257  -0.0741   0.6562   1.0000
   4.250   0.7959   0.03409   0.02438  -0.0746   0.6459   1.0000
   4.500   0.8198   0.03519   0.02554  -0.0738   0.6373   1.0000
   4.750   0.8401   0.03673   0.02719  -0.0737   0.6260   1.0000
   5.000   0.8607   0.03775   0.02834  -0.0725   0.6115   1.0000
   5.250   0.8829   0.03824   0.02890  -0.0702   0.5955   1.0000
   5.500   0.9066   0.03826   0.02898  -0.0674   0.5789   1.0000
   5.750   0.9271   0.03874   0.02957  -0.0652   0.5605   1.0000
   6.000   0.9475   0.03911   0.03012  -0.0630   0.5406   1.0000
   6.250   0.9717   0.03871   0.02981  -0.0599   0.5217   1.0000
   6.500   0.9977   0.03793   0.02911  -0.0565   0.5035   1.0000
   6.750   1.0178   0.03827   0.02967  -0.0544   0.4803   1.0000
   7.000   1.0468   0.03655   0.02799  -0.0503   0.4579   1.0000
   7.250   1.0701   0.03561   0.02715  -0.0469   0.4261   1.0000
   7.500   1.0941   0.03459   0.02613  -0.0435   0.3880   1.0000
   7.750   1.1174   0.03373   0.02515  -0.0402   0.3459   1.0000
   8.000   1.1380   0.03321   0.02440  -0.0373   0.3027   1.0000
   8.250   1.1554   0.03332   0.02429  -0.0350   0.2606   1.0000
   8.500   1.1706   0.03418   0.02487  -0.0329   0.2209   1.0000
   8.750   1.1840   0.03610   0.02638  -0.0309   0.1845   1.0000
   9.000   1.1995   0.03913   0.02913  -0.0293   0.1565   1.0000
   9.250   1.2157   0.04254   0.03260  -0.0279   0.1378   1.0000
   9.500   1.2344   0.04614   0.03623  -0.0268   0.1257   1.0000
   9.750   1.2455   0.04951   0.04007  -0.0255   0.1170   1.0000
  10.000   1.2573   0.05349   0.04428  -0.0245   0.1108   1.0000
  10.250   1.2589   0.05796   0.04935  -0.0232   0.1083   1.0000
  10.500   1.2560   0.06265   0.05452  -0.0222   0.1066   1.0000
  10.750   1.2469   0.06749   0.05977  -0.0215   0.1056   1.0000
  11.000   1.2287   0.07269   0.06532  -0.0213   0.1058   1.0000
  11.250   1.2008   0.07822   0.07110  -0.0218   0.1072   1.0000
  11.500   1.1713   0.08488   0.07793  -0.0246   0.1090   1.0000
  11.750   1.1439   0.09258   0.08574  -0.0289   0.1107   1.0000
  12.000   1.1214   0.10073   0.09394  -0.0332   0.1121   1.0000
<< Back to GOE 393 AIRFOIL (goe393-il)

Polar data table (+)

Polar graphs


<< Back to GOE 393 AIRFOIL (goe393-il)