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

GOE 147 (MVA H.6) AIRFOIL (goe147-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 147 (MVA H.6) AIRFOIL (goe147-il)
Reynolds number: 1,000,000
Max Cl/Cd: 102.76 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe147-il-1000000.txt
Download as CSV file: xf-goe147-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 147 (MVA H.6) AIRFOIL                       
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.000  -0.5758   0.15053   0.14890   0.0382   1.0000   0.0086
 -11.750  -0.5696   0.14710   0.14549   0.0369   1.0000   0.0088
  -7.750  -0.5611   0.02463   0.02184  -0.0485   1.0000   0.0087
  -7.500  -0.5365   0.02174   0.01863  -0.0496   1.0000   0.0089
  -7.250  -0.5111   0.01964   0.01626  -0.0502   1.0000   0.0091
  -7.000  -0.4851   0.01778   0.01415  -0.0505   1.0000   0.0093
  -6.750  -0.4585   0.01627   0.01241  -0.0507   1.0000   0.0096
  -6.500  -0.4314   0.01500   0.01094  -0.0507   1.0000   0.0099
  -6.250  -0.4039   0.01418   0.00998  -0.0507   1.0000   0.0104
  -6.000  -0.3761   0.01361   0.00931  -0.0506   1.0000   0.0107
  -5.750  -0.3487   0.01206   0.00753  -0.0507   1.0000   0.0114
  -5.500  -0.3210   0.01118   0.00656  -0.0507   1.0000   0.0120
  -5.250  -0.2931   0.01071   0.00604  -0.0506   1.0000   0.0127
  -5.000  -0.2652   0.01021   0.00549  -0.0505   1.0000   0.0133
  -4.750  -0.2361   0.00977   0.00499  -0.0506   0.9973   0.0140
  -4.500  -0.1969   0.00941   0.00455  -0.0528   0.9530   0.0148
  -4.250  -0.1729   0.00893   0.00377  -0.0514   0.8837   0.0162
  -4.000  -0.1475   0.00882   0.00345  -0.0505   0.8315   0.0177
  -3.750  -0.1204   0.00871   0.00315  -0.0501   0.7904   0.0192
  -3.500  -0.0926   0.00856   0.00283  -0.0498   0.7550   0.0205
  -3.250  -0.0645   0.00827   0.00237  -0.0496   0.7211   0.0246
  -3.000  -0.0363   0.00814   0.00208  -0.0494   0.6877   0.0294
  -2.750  -0.0083   0.00778   0.00183  -0.0494   0.6565   0.0794
  -2.500   0.0201   0.00805   0.00202  -0.0492   0.6303   0.0907
  -2.250   0.0485   0.00811   0.00200  -0.0491   0.6072   0.0952
  -2.000   0.0769   0.00823   0.00204  -0.0490   0.5860   0.0991
  -1.750   0.1054   0.00833   0.00207  -0.0489   0.5685   0.1022
  -1.500   0.1340   0.00843   0.00209  -0.0488   0.5546   0.1042
  -1.250   0.1625   0.00845   0.00205  -0.0488   0.5411   0.1056
  -0.750   0.2194   0.00821   0.00172  -0.0488   0.5142   0.1112
  -0.500   0.2479   0.00818   0.00164  -0.0487   0.4993   0.1124
  -0.250   0.2765   0.00816   0.00156  -0.0487   0.4817   0.1135
   0.000   0.3050   0.00815   0.00148  -0.0487   0.4637   0.1145
   0.250   0.3335   0.00817   0.00143  -0.0486   0.4426   0.1159
   0.500   0.3618   0.00824   0.00139  -0.0486   0.4139   0.1169
   0.750   0.3900   0.00837   0.00137  -0.0486   0.3761   0.1176
   1.000   0.4181   0.00853   0.00139  -0.0486   0.3434   0.1182
   1.250   0.4463   0.00865   0.00142  -0.0486   0.3223   0.1187
   1.500   0.4747   0.00872   0.00143  -0.0486   0.3068   0.1199
   1.750   0.5030   0.00879   0.00145  -0.0486   0.2940   0.1218
   2.000   0.5314   0.00884   0.00149  -0.0485   0.2837   0.1241
   2.250   0.5597   0.00890   0.00154  -0.0485   0.2744   0.1269
   2.500   0.5880   0.00899   0.00160  -0.0485   0.2651   0.1302
   2.750   0.6164   0.00903   0.00167  -0.0485   0.2575   0.1394
   3.000   0.6447   0.00904   0.00179  -0.0486   0.2496   0.1928
   3.250   0.6730   0.00907   0.00190  -0.0486   0.2426   0.2290
   3.500   0.7013   0.00907   0.00203  -0.0487   0.2353   0.2912
   3.750   0.7233   0.00761   0.00219  -0.0477   0.2293   1.0000
   4.000   0.7514   0.00778   0.00230  -0.0477   0.2216   1.0000
   4.250   0.7795   0.00791   0.00241  -0.0476   0.2148   1.0000
   4.500   0.8075   0.00808   0.00255  -0.0476   0.2067   1.0000
   4.750   0.8355   0.00825   0.00269  -0.0476   0.1977   1.0000
   5.000   0.8633   0.00843   0.00284  -0.0475   0.1874   1.0000
   5.250   0.8909   0.00867   0.00301  -0.0475   0.1723   1.0000
   5.500   0.9181   0.00898   0.00322  -0.0474   0.1547   1.0000
   5.750   0.9452   0.00931   0.00346  -0.0473   0.1405   1.0000
   6.000   0.9726   0.00955   0.00368  -0.0473   0.1335   1.0000
   6.250   0.9997   0.00984   0.00394  -0.0472   0.1264   1.0000
   6.500   1.0272   0.01003   0.00415  -0.0471   0.1231   1.0000
   6.750   1.0542   0.01031   0.00441  -0.0470   0.1173   1.0000
   7.000   1.0813   0.01055   0.00466  -0.0469   0.1117   1.0000
   7.250   1.1080   0.01085   0.00494  -0.0468   0.1036   1.0000
   7.500   1.1338   0.01134   0.00529  -0.0466   0.0767   1.0000
   7.750   1.1546   0.01294   0.00656  -0.0460   0.0206   1.0000
   8.000   1.1802   0.01344   0.00709  -0.0457   0.0169   1.0000
   8.250   1.2049   0.01412   0.00781  -0.0453   0.0139   1.0000
   8.500   1.2305   0.01455   0.00831  -0.0451   0.0130   1.0000
   8.750   1.2555   0.01507   0.00888  -0.0447   0.0120   1.0000
   9.000   1.2795   0.01575   0.00961  -0.0443   0.0109   1.0000
   9.250   1.3012   0.01686   0.01087  -0.0436   0.0099   1.0000
   9.500   1.3253   0.01741   0.01147  -0.0432   0.0096   1.0000
   9.750   1.3489   0.01802   0.01216  -0.0428   0.0090   1.0000
  10.000   1.3719   0.01867   0.01287  -0.0423   0.0085   1.0000
  10.250   1.3942   0.01940   0.01365  -0.0417   0.0080   1.0000
  10.500   1.4146   0.02035   0.01468  -0.0410   0.0076   1.0000
  10.750   1.4268   0.02233   0.01685  -0.0396   0.0071   1.0000
  11.000   1.4435   0.02356   0.01819  -0.0385   0.0069   1.0000
  11.250   1.4618   0.02451   0.01922  -0.0376   0.0067   1.0000
  11.500   1.4778   0.02562   0.02045  -0.0366   0.0065   1.0000
  11.750   1.4913   0.02691   0.02184  -0.0353   0.0063   1.0000
  12.000   1.5024   0.02830   0.02334  -0.0339   0.0061   1.0000
  12.250   1.5101   0.02983   0.02498  -0.0322   0.0059   1.0000
  12.500   1.5101   0.03153   0.02679  -0.0298   0.0058   1.0000
  12.750   1.5091   0.03370   0.02907  -0.0285   0.0057   1.0000
  13.000   1.5063   0.03667   0.03217  -0.0287   0.0056   1.0000
  13.250   1.5035   0.04030   0.03593  -0.0301   0.0055   1.0000
  13.500   1.5004   0.04446   0.04022  -0.0322   0.0054   1.0000
  13.750   1.4949   0.04922   0.04511  -0.0348   0.0054   1.0000
  14.000   1.4865   0.05447   0.05047  -0.0375   0.0053   1.0000
  14.250   1.4744   0.06017   0.05630  -0.0401   0.0053   1.0000
  14.500   1.4602   0.06621   0.06245  -0.0428   0.0052   1.0000
  14.750   1.4445   0.07245   0.06880  -0.0455   0.0052   1.0000
  15.000   1.4282   0.07885   0.07532  -0.0481   0.0052   1.0000
  15.250   1.4121   0.08522   0.08179  -0.0507   0.0052   1.0000
<< Back to GOE 147 (MVA H.6) AIRFOIL (goe147-il)

Polar data table (+)

Polar graphs


<< Back to GOE 147 (MVA H.6) AIRFOIL (goe147-il)