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

GOE 804 (EA 8) AIRFOIL (goe804-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: GOE 804 (EA 8) AIRFOIL (goe804-il)
Reynolds number: 1,000,000
Max Cl/Cd: 147.02 at α=0°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe804-il-1000000-n5.txt
Download as CSV file: xf-goe804-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 804 (EA 8) 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.0296   0.09451   0.09269  -0.1075   0.9610   0.0047
 -10.750  -0.0218   0.09123   0.08942  -0.1089   0.9596   0.0051
 -10.500  -0.0134   0.08791   0.08609  -0.1105   0.9584   0.0051
 -10.250  -0.0048   0.08397   0.08216  -0.1125   0.9575   0.0055
 -10.000   0.0047   0.08039   0.07857  -0.1146   0.9566   0.0055
  -9.500   0.0200   0.07373   0.07191  -0.1174   0.9519   0.0055
  -9.000   0.0362   0.06585   0.06402  -0.1212   0.9471   0.0038
  -8.750   0.0477   0.06194   0.06010  -0.1241   0.9454   0.0038
  -8.500  -0.0285   0.08111   0.07922  -0.1196   0.9523   0.0041
  -8.250  -0.0144   0.07844   0.07654  -0.1223   0.9484   0.0056
  -8.000   0.0007   0.07528   0.07337  -0.1254   0.9459   0.0036
  -7.750   0.0942   0.04642   0.04455  -0.1364   0.9343   0.0041
  -7.500   0.1116   0.04464   0.04275  -0.1388   0.9313   0.0047
  -7.000   0.0764   0.06069   0.05873  -0.1452   0.9314   0.0043
  -6.500   0.1824   0.01312   0.00957  -0.2173   0.9122   0.0042
  -6.250   0.2125   0.01178   0.00797  -0.2189   0.9104   0.0044
  -6.000   0.2424   0.01085   0.00684  -0.2200   0.9085   0.0046
  -5.750   0.2727   0.01004   0.00584  -0.2211   0.9065   0.0050
  -5.500   0.3032   0.00938   0.00500  -0.2221   0.9046   0.0053
  -5.250   0.3339   0.00888   0.00434  -0.2230   0.9024   0.0056
  -5.000   0.3650   0.00838   0.00373  -0.2241   0.9002   0.0063
  -4.750   0.3958   0.00808   0.00334  -0.2250   0.8980   0.0071
  -4.500   0.4243   0.00778   0.00296  -0.2253   0.8958   0.0077
  -4.250   0.4530   0.00744   0.00256  -0.2257   0.8930   0.0087
  -4.000   0.4817   0.00722   0.00229  -0.2260   0.8897   0.0100
  -3.750   0.5116   0.00699   0.00200  -0.2266   0.8852   0.0122
  -3.500   0.5400   0.00681   0.00177  -0.2268   0.8802   0.0150
  -3.250   0.5682   0.00664   0.00158  -0.2270   0.8756   0.0205
  -3.000   0.5972   0.00650   0.00148  -0.2274   0.8727   0.0326
  -2.750   0.6261   0.00644   0.00141  -0.2277   0.8701   0.0404
  -2.500   0.6543   0.00640   0.00136  -0.2279   0.8673   0.0458
  -2.250   0.6816   0.00637   0.00133  -0.2278   0.8638   0.0489
  -2.000   0.7097   0.00633   0.00126  -0.2280   0.8604   0.0511
  -1.750   0.7382   0.00625   0.00116  -0.2282   0.8558   0.0552
  -1.500   0.7655   0.00620   0.00110  -0.2282   0.8492   0.0575
  -1.250   0.7930   0.00617   0.00104  -0.2281   0.8426   0.0589
  -1.000   0.8205   0.00615   0.00099  -0.2281   0.8364   0.0602
  -0.750   0.8469   0.00615   0.00095  -0.2278   0.8247   0.0614
  -0.500   0.8726   0.00618   0.00092  -0.2274   0.8084   0.0646
  -0.250   0.8984   0.00621   0.00093  -0.2270   0.7922   0.0720
   0.000   0.9233   0.00628   0.00097  -0.2264   0.7719   0.0884
   0.250   0.9453   0.00649   0.00104  -0.2251   0.7391   0.0987
   0.500   0.9656   0.00679   0.00116  -0.2235   0.7012   0.1048
   0.750   0.9866   0.00708   0.00132  -0.2221   0.6666   0.1124
   1.000   1.0085   0.00736   0.00147  -0.2209   0.6371   0.1191
   1.250   1.0310   0.00761   0.00166  -0.2199   0.6074   0.1462
   1.500   1.0533   0.00789   0.00189  -0.2189   0.5719   0.2023
   1.750   1.0692   0.00857   0.00223  -0.2166   0.4893   0.2239
   2.000   1.0915   0.00899   0.00268  -0.2160   0.4269   0.4124
   2.250   1.1169   0.00919   0.00296  -0.2158   0.4017   0.5092
   2.500   1.1465   0.00911   0.00340  -0.2167   0.3742   0.8172
   2.750   1.1636   0.00982   0.00381  -0.2147   0.2961   0.8580
   3.000   1.1818   0.01043   0.00421  -0.2129   0.2417   0.8824
   3.250   1.2029   0.01071   0.00446  -0.2116   0.2239   0.9090
   3.500   1.2177   0.01080   0.00461  -0.2088   0.2115   0.9624
   3.750   1.2393   0.01125   0.00493  -0.2077   0.1788   1.0000
   4.000   1.2560   0.01218   0.00547  -0.2058   0.1061   1.0000
   4.250   1.2762   0.01281   0.00591  -0.2044   0.0678   1.0000
   4.500   1.2977   0.01326   0.00629  -0.2033   0.0527   1.0000
   4.750   1.3197   0.01362   0.00664  -0.2023   0.0417   1.0000
   5.000   1.3379   0.01434   0.00720  -0.2006   0.0152   1.0000
   5.250   1.3586   0.01481   0.00765  -0.1993   0.0101   1.0000
   5.500   1.3793   0.01524   0.00810  -0.1980   0.0079   1.0000
   5.750   1.3997   0.01571   0.00859  -0.1967   0.0065   1.0000
   6.000   1.4203   0.01615   0.00905  -0.1954   0.0055   1.0000
   6.250   1.4397   0.01669   0.00962  -0.1939   0.0047   1.0000
   6.500   1.4594   0.01719   0.01018  -0.1925   0.0043   1.0000
   6.750   1.4785   0.01773   0.01076  -0.1910   0.0038   1.0000
   7.000   1.4969   0.01837   0.01144  -0.1893   0.0035   1.0000
   7.250   1.5145   0.01907   0.01221  -0.1876   0.0032   1.0000
   7.500   1.5326   0.01970   0.01290  -0.1859   0.0029   1.0000
   7.750   1.5505   0.02034   0.01361  -0.1843   0.0027   1.0000
   8.000   1.5678   0.02106   0.01438  -0.1826   0.0025   1.0000
   8.250   1.5836   0.02191   0.01530  -0.1806   0.0023   1.0000
   8.500   1.5975   0.02298   0.01645  -0.1783   0.0022   1.0000
   8.750   1.6131   0.02386   0.01743  -0.1764   0.0021   1.0000
   9.000   1.6271   0.02492   0.01859  -0.1742   0.0020   1.0000
   9.250   1.6409   0.02598   0.01977  -0.1720   0.0019   1.0000
   9.500   1.6539   0.02715   0.02105  -0.1698   0.0018   1.0000
   9.750   1.6657   0.02844   0.02246  -0.1674   0.0017   1.0000
  10.000   1.6778   0.02969   0.02382  -0.1651   0.0017   1.0000
  10.250   1.6904   0.03089   0.02512  -0.1631   0.0016   1.0000
  10.500   1.7008   0.03234   0.02668  -0.1607   0.0015   1.0000
  10.750   1.7114   0.03376   0.02821  -0.1585   0.0015   1.0000
  11.000   1.7181   0.03567   0.03028  -0.1558   0.0014   1.0000
  11.250   1.7227   0.03791   0.03270  -0.1529   0.0013   1.0000
  11.500   1.7285   0.03995   0.03492  -0.1502   0.0013   1.0000
  11.750   1.7339   0.04203   0.03719  -0.1477   0.0012   1.0000
  12.000   1.7360   0.04461   0.03997  -0.1449   0.0012   1.0000
  12.250   1.7382   0.04712   0.04271  -0.1423   0.0012   1.0000
  12.500   1.7379   0.04996   0.04575  -0.1397   0.0011   1.0000
  12.750   1.7354   0.05309   0.04910  -0.1371   0.0011   1.0000
  13.000   1.7261   0.05729   0.05357  -0.1342   0.0011   1.0000
  13.250   1.7183   0.06125   0.05775  -0.1319   0.0011   1.0000
  13.500   1.7102   0.06538   0.06209  -0.1300   0.0010   1.0000
  13.750   1.6935   0.07095   0.06793  -0.1282   0.0010   1.0000
  14.000   1.6659   0.07867   0.07598  -0.1270   0.0010   1.0000
  14.250   1.6542   0.08406   0.08157  -0.1270   0.0010   1.0000
  14.500   1.6293   0.09215   0.08992  -0.1281   0.0010   1.0000
  14.750   1.6112   0.09951   0.09748  -0.1300   0.0010   1.0000
  15.000   1.5785   0.11067   0.10891  -0.1344   0.0010   1.0000
  15.250   1.5499   0.12202   0.12050  -0.1403   0.0010   1.0000
<< Back to GOE 804 (EA 8) AIRFOIL (goe804-il)

Polar data table (+)

Polar graphs


<< Back to GOE 804 (EA 8) AIRFOIL (goe804-il)