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

GOE 118 (MVA MK.7) AIRFOIL (goe118-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 118 (MVA MK.7) AIRFOIL (goe118-il)
Reynolds number: 500,000
Max Cl/Cd: 108.94 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe118-il-500000-n5.txt
Download as CSV file: xf-goe118-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 118 (MVA MK.7) 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.2245   0.09991   0.09755  -0.0487   0.8845   0.0061
  -8.500  -0.2179   0.09701   0.09456  -0.0500   0.8623   0.0062
  -8.250  -0.2109   0.09407   0.09154  -0.0515   0.8430   0.0062
  -8.000  -0.2035   0.09101   0.08842  -0.0532   0.8253   0.0063
  -7.750  -0.1958   0.08787   0.08522  -0.0550   0.8096   0.0063
  -7.500  -0.1876   0.08468   0.08195  -0.0569   0.7953   0.0063
  -7.250  -0.1793   0.08152   0.07874  -0.0591   0.7825   0.0063
  -7.000  -0.1724   0.07689   0.07408  -0.0627   0.7718   0.0065
  -6.500  -0.1443   0.06984   0.06693  -0.0700   0.7524   0.0068
  -6.250  -0.1267   0.06631   0.06335  -0.0743   0.7439   0.0069
  -6.000  -0.1064   0.06260   0.05958  -0.0792   0.7356   0.0071
  -5.750  -0.0834   0.05864   0.05556  -0.0847   0.7282   0.0073
  -4.750   0.0400   0.03872   0.03515  -0.1107   0.7022   0.0056
  -4.500   0.0747   0.03298   0.02917  -0.1164   0.6961   0.0050
  -4.250   0.1142   0.02301   0.01858  -0.1222   0.6916   0.0043
  -4.000   0.1471   0.01769   0.01242  -0.1242   0.6864   0.0040
  -3.750   0.1762   0.01560   0.00986  -0.1248   0.6802   0.0041
  -3.500   0.2049   0.01424   0.00813  -0.1251   0.6743   0.0044
  -3.250   0.2336   0.01314   0.00674  -0.1253   0.6678   0.0049
  -3.000   0.2619   0.01274   0.00612  -0.1253   0.6618   0.0061
  -2.750   0.2905   0.01164   0.00482  -0.1255   0.6555   0.0066
  -2.500   0.3187   0.01080   0.00381  -0.1256   0.6492   0.0071
  -2.250   0.3471   0.01032   0.00325  -0.1258   0.6430   0.0081
  -2.000   0.3753   0.01003   0.00285  -0.1259   0.6362   0.0099
  -1.750   0.4037   0.00972   0.00238  -0.1260   0.6295   0.0129
  -1.500   0.4318   0.00950   0.00203  -0.1261   0.6212   0.0186
  -1.250   0.4598   0.00920   0.00180  -0.1263   0.6124   0.0710
  -1.000   0.4875   0.00923   0.00178  -0.1264   0.6042   0.0898
  -0.750   0.5154   0.00923   0.00175  -0.1266   0.5967   0.0987
  -0.500   0.5431   0.00925   0.00171  -0.1267   0.5903   0.1037
  -0.250   0.5711   0.00923   0.00164  -0.1269   0.5838   0.1052
   0.000   0.5987   0.00925   0.00158  -0.1269   0.5765   0.1067
   0.250   0.6264   0.00924   0.00155  -0.1271   0.5663   0.1106
   0.500   0.6539   0.00926   0.00154  -0.1272   0.5568   0.1154
   1.000   0.7090   0.00932   0.00158  -0.1274   0.5404   0.1324
   1.250   0.7364   0.00934   0.00162  -0.1275   0.5326   0.1512
   1.500   0.7639   0.00924   0.00170  -0.1278   0.5239   0.2278
   1.750   0.7905   0.00860   0.00194  -0.1282   0.5154   0.6005
   2.250   0.8409   0.00803   0.00208  -0.1272   0.4983   1.0000
   2.500   0.8679   0.00818   0.00220  -0.1273   0.4908   1.0000
   2.750   0.8949   0.00832   0.00232  -0.1273   0.4794   1.0000
   3.250   0.9478   0.00870   0.00261  -0.1272   0.4496   1.0000
   3.500   0.9714   0.00916   0.00287  -0.1267   0.4051   1.0000
   3.750   0.9911   0.01006   0.00334  -0.1257   0.3258   1.0000
   4.000   0.9997   0.01224   0.00456  -0.1232   0.1394   1.0000
   4.250   1.0165   0.01349   0.00544  -0.1217   0.0219   1.0000
   4.500   1.0405   0.01391   0.00592  -0.1211   0.0129   1.0000
   4.750   1.0630   0.01450   0.00661  -0.1203   0.0084   1.0000
   5.000   1.0855   0.01504   0.00725  -0.1195   0.0073   1.0000
   5.250   1.1080   0.01552   0.00774  -0.1188   0.0057   1.0000
   5.500   1.1264   0.01641   0.00874  -0.1174   0.0049   1.0000
   5.750   1.1439   0.01731   0.00978  -0.1158   0.0044   1.0000
   6.000   1.1585   0.01839   0.01098  -0.1137   0.0041   1.0000
   6.250   1.1699   0.01962   0.01232  -0.1111   0.0039   1.0000
   6.500   1.1786   0.02095   0.01376  -0.1082   0.0037   1.0000
   6.750   1.1871   0.02204   0.01492  -0.1052   0.0035   1.0000
   7.000   1.1940   0.02325   0.01617  -0.1024   0.0031   1.0000
   7.250   1.2014   0.02462   0.01767  -0.0994   0.0027   1.0000
   7.500   1.2074   0.02647   0.01962  -0.0962   0.0025   1.0000
   7.750   1.2219   0.02845   0.02168  -0.0938   0.0023   1.0000
   8.000   1.2552   0.03092   0.02423  -0.0937   0.0021   1.0000
   8.250   1.3017   0.03450   0.02802  -0.0958   0.0022   1.0000
   8.500   1.3316   0.03796   0.03171  -0.0954   0.0023   1.0000
   8.750   1.3519   0.04121   0.03521  -0.0939   0.0025   1.0000
   9.000   1.3660   0.04441   0.03865  -0.0919   0.0026   1.0000
   9.250   1.3738   0.04767   0.04214  -0.0896   0.0028   1.0000
  13.750   1.1464   0.12680   0.12451  -0.0723   0.0039   1.0000
  14.000   1.1293   0.13473   0.13255  -0.0771   0.0039   1.0000
  14.250   1.1134   0.14330   0.14122  -0.0826   0.0040   1.0000
<< Back to GOE 118 (MVA MK.7) AIRFOIL (goe118-il)

Polar data table (+)

Polar graphs


<< Back to GOE 118 (MVA MK.7) AIRFOIL (goe118-il)