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

GOE 775 AIRFOIL (goe775-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 775 AIRFOIL (goe775-il)
Reynolds number: 50,000
Max Cl/Cd: 18.85 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe775-il-50000.txt
Download as CSV file: xf-goe775-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 775 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
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.2941   0.11861   0.10934  -0.0089   1.0000   0.4776
  -9.500  -0.2700   0.11551   0.10623  -0.0093   1.0000   0.4889
  -9.250  -0.3055   0.11536   0.10611  -0.0069   1.0000   0.4958
  -9.000  -0.2573   0.11013   0.10087  -0.0089   1.0000   0.5048
  -8.750  -0.2718   0.10903   0.09980  -0.0072   1.0000   0.5156
  -8.500  -0.2788   0.10260   0.09335  -0.0096   1.0000   0.4824
  -8.250  -0.3976   0.08671   0.07741  -0.0149   1.0000   0.4026
  -8.000  -0.4322   0.08147   0.07222  -0.0136   1.0000   0.4080
  -7.750  -0.6257   0.06372   0.05465  -0.0082   1.0000   0.4114
  -7.500  -0.8111   0.05270   0.04363   0.0100   1.0000   0.4114
  -7.250  -0.5891   0.06355   0.05453  -0.0036   1.0000   0.4300
  -7.000  -0.5999   0.06164   0.05268   0.0004   1.0000   0.4380
  -6.750  -0.7164   0.05460   0.04573   0.0131   1.0000   0.4424
  -6.500  -0.7982   0.04916   0.04019   0.0237   1.0000   0.4493
  -6.250  -0.8066   0.04763   0.03865   0.0281   1.0000   0.4586
  -6.000  -0.8015   0.04704   0.03807   0.0313   1.0000   0.4693
  -5.750  -0.8102   0.04548   0.03648   0.0356   1.0000   0.4794
  -5.500  -0.8019   0.04510   0.03612   0.0384   1.0000   0.4908
  -5.250  -0.7984   0.04430   0.03532   0.0415   1.0000   0.5017
  -5.000  -0.7991   0.04329   0.03426   0.0449   1.0000   0.5142
  -4.750  -0.7825   0.04340   0.03445   0.0469   1.0000   0.5251
  -4.500  -0.7927   0.04157   0.03247   0.0506   1.0000   0.5393
  -4.250  -0.7660   0.04251   0.03358   0.0518   1.0000   0.5497
  -4.000  -0.7665   0.04130   0.03230   0.0547   1.0000   0.5635
  -3.750  -0.7508   0.04146   0.03252   0.0565   1.0000   0.5759
  -3.500  -0.7417   0.04100   0.03207   0.0587   1.0000   0.5888
  -3.250  -0.7393   0.03998   0.03095   0.0610   1.0000   0.6047
  -3.000  -0.7184   0.04066   0.03176   0.0625   1.0000   0.6158
  -2.750  -0.6825   0.04092   0.03201   0.0599   0.9922   0.6331
  -2.500  -0.6320   0.04154   0.03262   0.0551   0.9781   0.6526
  -2.250  -0.5891   0.04185   0.03288   0.0514   0.9635   0.6720
  -2.000  -0.5325   0.04336   0.03447   0.0474   0.9486   0.6866
  -1.750  -0.4872   0.04407   0.03520   0.0443   0.9334   0.7029
  -1.500  -0.4485   0.04449   0.03560   0.0420   0.9188   0.7205
  -1.250  -0.4045   0.04497   0.03605   0.0389   0.9046   0.7392
  -1.000  -0.3166   0.04710   0.03822   0.0305   0.8895   0.7521
  -0.750  -0.2767   0.04770   0.03882   0.0285   0.8735   0.7673
  -0.500  -0.2367   0.04806   0.03916   0.0261   0.8586   0.7845
  -0.250  -0.1100   0.04948   0.04056   0.0112   0.8453   0.7987
   0.000  -0.0265   0.05005   0.04115   0.0021   0.8285   0.8118
   0.250   0.0137   0.05006   0.04116  -0.0005   0.8124   0.8280
   0.500   0.1178   0.04942   0.04050  -0.0126   0.7998   0.8445
   0.750   0.2298   0.04822   0.03932  -0.0255   0.7849   0.8580
   1.000   0.2692   0.04787   0.03899  -0.0278   0.7679   0.8730
   1.250   0.3131   0.04717   0.03828  -0.0304   0.7532   0.8893
   1.500   0.4006   0.04513   0.03621  -0.0388   0.7398   0.9041
   1.750   0.4445   0.04465   0.03576  -0.0419   0.7213   0.9185
   2.000   0.4841   0.04420   0.03532  -0.0443   0.7039   0.9332
   2.250   0.5321   0.04333   0.03443  -0.0476   0.6882   0.9484
   2.500   0.5859   0.04201   0.03306  -0.0515   0.6728   0.9634
   2.750   0.6293   0.04169   0.03278  -0.0552   0.6536   0.9780
   3.000   0.6806   0.04103   0.03213  -0.0602   0.6343   0.9921
   3.250   0.7176   0.04068   0.03179  -0.0628   0.6173   1.0000
   3.500   0.7368   0.04013   0.03113  -0.0612   0.6057   1.0000
   3.750   0.7397   0.04112   0.03219  -0.0589   0.5901   1.0000
   4.000   0.7518   0.04132   0.03235  -0.0570   0.5778   1.0000
   4.250   0.7633   0.04149   0.03251  -0.0549   0.5647   1.0000
   4.500   0.7649   0.04254   0.03361  -0.0521   0.5513   1.0000
   4.750   0.7884   0.04182   0.03275  -0.0507   0.5404   1.0000
   5.000   0.7792   0.04357   0.03463  -0.0470   0.5266   1.0000
   5.250   0.8050   0.04281   0.03372  -0.0457   0.5161   1.0000
   5.500   0.7924   0.04463   0.03568  -0.0414   0.5032   1.0000
   5.750   0.8067   0.04472   0.03570  -0.0392   0.4925   1.0000
   6.000   0.8002   0.04604   0.03707  -0.0351   0.4810   1.0000
   6.250   0.8074   0.04661   0.03761  -0.0322   0.4708   1.0000
   6.500   0.7902   0.04851   0.03957  -0.0271   0.4604   1.0000
   6.750   0.8073   0.04857   0.03956  -0.0249   0.4505   1.0000
   7.000   0.7074   0.05503   0.04614  -0.0127   0.4442   1.0000
   7.250   0.6283   0.05997   0.05103  -0.0027   0.4383   1.0000
   7.500   0.6314   0.06091   0.05192   0.0007   0.4298   1.0000
   7.750   0.8257   0.05187   0.04275  -0.0117   0.4120   1.0000
   8.000   0.6591   0.06164   0.05258   0.0058   0.4114   1.0000
   8.250   0.4385   0.08085   0.07158   0.0134   0.4087   1.0000
   8.500   0.3993   0.08647   0.07715   0.0146   0.4035   1.0000
   8.750   0.2858   0.10634   0.09713   0.0061   0.5168   1.0000
   9.000   0.2752   0.10897   0.09973   0.0069   0.5160   1.0000
   9.250   0.2575   0.10999   0.10072   0.0087   0.5058   1.0000
   9.500   0.3088   0.11527   0.10602   0.0066   0.4960   1.0000
   9.750   0.2727   0.11551   0.10622   0.0089   0.4902   1.0000
  10.000   0.2916   0.11825   0.10896   0.0088   0.4784   1.0000
<< Back to GOE 775 AIRFOIL (goe775-il)

Polar data table (+)

Polar graphs


<< Back to GOE 775 AIRFOIL (goe775-il)