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

GOE 233 (MVA CA4) AIRFOIL (goe233-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 233 (MVA CA4) AIRFOIL (goe233-il)
Reynolds number: 500,000
Max Cl/Cd: 120.74 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe233-il-500000.txt
Download as CSV file: xf-goe233-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 233 (MVA CA4) AIRFOIL                       
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.0421   0.09284   0.09049  -0.0791   0.9333   0.0316
 -10.000  -0.0545   0.08687   0.08450  -0.0829   0.9250   0.0343
  -9.750  -0.0510   0.08306   0.08068  -0.0830   0.9175   0.0347
  -9.500  -0.0401   0.08121   0.07880  -0.0824   0.9094   0.0352
  -9.250  -0.0294   0.07928   0.07684  -0.0822   0.9020   0.0358
  -9.000  -0.0211   0.07686   0.07440  -0.0824   0.8937   0.0366
  -8.750  -0.0156   0.07386   0.07136  -0.0832   0.8863   0.0380
  -8.500  -0.0280   0.06671   0.06420  -0.0875   0.8778   0.0409
  -8.250  -0.0143   0.06545   0.06290  -0.0867   0.8706   0.0416
  -8.000  -0.0014   0.06382   0.06126  -0.0867   0.8627   0.0426
  -7.750   0.0063   0.06116   0.05854  -0.0876   0.8556   0.0447
  -7.500   0.0059   0.02172   0.01735  -0.1911   0.8682   0.0349
  -7.250   0.0417   0.01977   0.01513  -0.1947   0.8616   0.0349
  -7.000   0.0739   0.01810   0.01329  -0.1969   0.8559   0.0355
  -6.750   0.1064   0.01689   0.01194  -0.1987   0.8493   0.0357
  -6.500   0.1380   0.01591   0.01081  -0.1999   0.8427   0.0358
  -6.250   0.1692   0.01510   0.00987  -0.2009   0.8363   0.0360
  -6.000   0.2003   0.01439   0.00909  -0.2018   0.8290   0.0364
  -5.750   0.2311   0.01382   0.00841  -0.2025   0.8227   0.0368
  -5.500   0.2622   0.01330   0.00783  -0.2034   0.8154   0.0376
  -5.250   0.2931   0.01285   0.00729  -0.2040   0.8082   0.0382
  -5.000   0.3243   0.01243   0.00680  -0.2047   0.8004   0.0386
  -4.750   0.3553   0.01205   0.00632  -0.2054   0.7918   0.0391
  -4.500   0.3865   0.01169   0.00588  -0.2061   0.7834   0.0395
  -4.250   0.4176   0.01135   0.00545  -0.2068   0.7751   0.0400
  -4.000   0.4489   0.01104   0.00504  -0.2076   0.7676   0.0405
  -3.750   0.4807   0.01062   0.00453  -0.2085   0.7596   0.0419
  -3.500   0.5117   0.01034   0.00414  -0.2091   0.7524   0.0437
  -3.250   0.5422   0.01012   0.00384  -0.2096   0.7446   0.0454
  -3.000   0.5723   0.00997   0.00360  -0.2099   0.7376   0.0482
  -2.750   0.6030   0.00971   0.00339  -0.2105   0.7297   0.0677
  -2.500   0.6331   0.00953   0.00318  -0.2109   0.7227   0.0914
  -2.250   0.6639   0.00922   0.00301  -0.2116   0.7156   0.1495
  -2.000   0.6936   0.00906   0.00287  -0.2120   0.7088   0.1810
  -1.750   0.7232   0.00893   0.00287  -0.2123   0.7026   0.2388
  -1.500   0.7519   0.00898   0.00293  -0.2123   0.6952   0.2644
  -1.250   0.7803   0.00910   0.00298  -0.2121   0.6880   0.2792
  -1.000   0.8087   0.00920   0.00309  -0.2120   0.6799   0.2896
  -0.750   0.8369   0.00935   0.00315  -0.2118   0.6731   0.2990
  -0.500   0.8652   0.00951   0.00332  -0.2116   0.6669   0.3065
  -0.250   0.8935   0.00966   0.00341  -0.2115   0.6606   0.3135
   0.000   0.9212   0.00988   0.00363  -0.2112   0.6545   0.3191
   0.250   0.9493   0.01006   0.00381  -0.2110   0.6474   0.3258
   0.500   0.9771   0.01022   0.00391  -0.2108   0.6406   0.3308
   0.750   1.0050   0.01037   0.00409  -0.2105   0.6345   0.3342
   1.000   1.0330   0.01049   0.00421  -0.2104   0.6284   0.3375
   1.250   1.0609   0.01064   0.00430  -0.2102   0.6227   0.3413
   1.500   1.0892   0.01072   0.00437  -0.2101   0.6167   0.3444
   1.750   1.1170   0.01078   0.00444  -0.2100   0.6100   0.3476
   2.000   1.1444   0.01095   0.00460  -0.2097   0.6037   0.3505
   2.250   1.1721   0.01102   0.00472  -0.2096   0.5965   0.3535
   2.500   1.1995   0.01115   0.00479  -0.2094   0.5893   0.3561
   2.750   1.2273   0.01120   0.00487  -0.2092   0.5819   0.3589
   3.000   1.2547   0.01132   0.00495  -0.2090   0.5747   0.3622
   3.250   1.2818   0.01140   0.00508  -0.2088   0.5675   0.3651
   3.500   1.3088   0.01150   0.00520  -0.2085   0.5588   0.3675
   3.750   1.3359   0.01160   0.00531  -0.2083   0.5507   0.3697
   4.000   1.3628   0.01170   0.00541  -0.2080   0.5422   0.3720
   4.250   1.3896   0.01181   0.00553  -0.2078   0.5339   0.3743
   4.500   1.4160   0.01195   0.00564  -0.2074   0.5244   0.3764
   4.750   1.4426   0.01204   0.00576  -0.2071   0.5145   0.3784
   5.000   1.4683   0.01219   0.00592  -0.2067   0.5046   0.3804
   5.250   1.4936   0.01237   0.00610  -0.2062   0.4920   0.3827
   5.500   1.5185   0.01259   0.00632  -0.2056   0.4779   0.3852
   5.750   1.5430   0.01283   0.00654  -0.2049   0.4639   0.3878
   6.000   1.5669   0.01312   0.00679  -0.2042   0.4492   0.3902
   6.250   1.5901   0.01345   0.00707  -0.2033   0.4328   0.3923
   6.500   1.6121   0.01382   0.00739  -0.2023   0.4130   0.3944
   6.750   1.6328   0.01428   0.00777  -0.2010   0.3896   0.3965
   7.000   1.6528   0.01478   0.00820  -0.1997   0.3630   0.3987
   7.250   1.6703   0.01544   0.00872  -0.1980   0.3299   0.4011
   7.500   1.6839   0.01634   0.00939  -0.1957   0.2903   0.4037
   7.750   1.6968   0.01724   0.01012  -0.1932   0.2636   0.4062
   8.000   1.7110   0.01800   0.01080  -0.1910   0.2469   0.4082
   8.500   1.7362   0.01927   0.01207  -0.1858   0.2244   0.4125
   8.750   1.7478   0.01997   0.01277  -0.1832   0.2149   0.4147
   9.000   1.7599   0.02074   0.01354  -0.1807   0.2047   0.4171
   9.250   1.7737   0.02143   0.01426  -0.1787   0.1941   0.4197
   9.500   1.7851   0.02230   0.01513  -0.1763   0.1807   0.4222
   9.750   1.7920   0.02351   0.01622  -0.1735   0.1571   0.4245
  10.000   1.7842   0.02578   0.01818  -0.1691   0.1089   0.4264
  10.250   1.7727   0.02853   0.02071  -0.1646   0.0764   0.4279
  10.500   1.7702   0.03074   0.02287  -0.1614   0.0616   0.4299
  10.750   1.7722   0.03269   0.02482  -0.1589   0.0538   0.4321
  11.000   1.7743   0.03470   0.02686  -0.1565   0.0487   0.4345
  11.250   1.7788   0.03656   0.02878  -0.1544   0.0453   0.4369
  11.500   1.7788   0.03889   0.03114  -0.1522   0.0422   0.4388
  11.750   1.7836   0.04083   0.03319  -0.1504   0.0401   0.4415
  12.000   1.7873   0.04291   0.03537  -0.1487   0.0382   0.4441
  12.250   1.7873   0.04541   0.03794  -0.1469   0.0365   0.4468
  12.500   1.7812   0.04864   0.04126  -0.1449   0.0347   0.4492
  12.750   1.7858   0.05083   0.04355  -0.1437   0.0335   0.4527
  13.000   1.7880   0.05334   0.04616  -0.1424   0.0322   0.4561
  13.250   1.7883   0.05612   0.04904  -0.1412   0.0309   0.4594
  13.500   1.7843   0.05948   0.05248  -0.1400   0.0298   0.4625
  13.750   1.7731   0.06383   0.05693  -0.1388   0.0287   0.4650
  14.000   1.7759   0.06652   0.05975  -0.1380   0.0280   0.4691
  14.250   1.7755   0.06966   0.06301  -0.1373   0.0271   0.4732
  14.500   1.7740   0.07300   0.06646  -0.1367   0.0263   0.4778
  14.750   1.7715   0.07651   0.07007  -0.1362   0.0255   0.4830
  15.000   1.7663   0.08043   0.07408  -0.1358   0.0248   0.4882
  15.250   1.7555   0.08515   0.07890  -0.1355   0.0241   0.4936
<< Back to GOE 233 (MVA CA4) AIRFOIL (goe233-il)

Polar data table (+)

Polar graphs


<< Back to GOE 233 (MVA CA4) AIRFOIL (goe233-il)