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

GOE 380 AIRFOIL (goe380-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 380 AIRFOIL (goe380-il)
Reynolds number: 200,000
Max Cl/Cd: 78.01 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe380-il-200000-n5.txt
Download as CSV file: xf-goe380-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 380 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.2619   0.10044   0.09720  -0.0254   1.0000   0.0136
  -7.750  -0.2634   0.09889   0.09571  -0.0246   1.0000   0.0139
  -7.500  -0.2680   0.09742   0.09431  -0.0231   1.0000   0.0140
  -7.250  -0.2572   0.09520   0.09213  -0.0268   0.9951   0.0143
  -7.000  -0.2345   0.09177   0.08870  -0.0336   0.9875   0.0145
  -6.750  -0.2096   0.08822   0.08514  -0.0409   0.9797   0.0147
  -6.500  -0.1856   0.08452   0.08143  -0.0468   0.9725   0.0148
  -5.500  -0.0799   0.06793   0.06472  -0.0690   0.9440   0.0153
  -5.250  -0.0625   0.06257   0.05933  -0.0718   0.9368   0.0160
  -5.000  -0.0438   0.05913   0.05588  -0.0732   0.9284   0.0176
  -4.750  -0.0171   0.05585   0.05254  -0.0770   0.9183   0.0185
  -4.500   0.0126   0.05245   0.04906  -0.0813   0.9090   0.0197
  -4.250   0.0436   0.04910   0.04559  -0.0853   0.8988   0.0210
  -4.000   0.0737   0.04597   0.04234  -0.0885   0.8865   0.0227
  -3.750   0.1153   0.04325   0.03936  -0.0930   0.8733   0.0260
  -3.500   0.1498   0.04046   0.03631  -0.0955   0.8594   0.0264
  -3.000   0.2065   0.03482   0.03022  -0.0974   0.8266   0.0265
  -2.750   0.2278   0.03049   0.02570  -0.0972   0.8095   0.0142
  -2.500   0.2542   0.02773   0.02264  -0.0973   0.7923   0.0135
  -2.250   0.2816   0.02527   0.01987  -0.0970   0.7756   0.0138
  -2.000   0.3093   0.02323   0.01750  -0.0964   0.7596   0.0147
  -1.750   0.3361   0.02095   0.01485  -0.0956   0.7432   0.0150
  -1.500   0.3622   0.01826   0.01168  -0.0946   0.7265   0.0148
  -1.250   0.3885   0.01616   0.00905  -0.0936   0.7086   0.0150
  -1.000   0.4146   0.01422   0.00657  -0.0926   0.6908   0.0156
  -0.750   0.4399   0.01312   0.00513  -0.0917   0.6723   0.0191
  -0.500   0.4649   0.01253   0.00430  -0.0908   0.6534   0.0219
  -0.250   0.4898   0.01212   0.00367  -0.0899   0.6353   0.0257
   0.000   0.5145   0.01188   0.00340  -0.0891   0.6185   0.0434
   0.250   0.5406   0.01210   0.00350  -0.0885   0.6032   0.0722
   0.500   0.5666   0.01235   0.00358  -0.0881   0.5890   0.0875
   0.750   0.5918   0.01235   0.00345  -0.0875   0.5758   0.0955
   1.000   0.6166   0.01229   0.00329  -0.0869   0.5634   0.1017
   1.250   0.6414   0.01226   0.00317  -0.0862   0.5519   0.1077
   1.500   0.6664   0.01225   0.00307  -0.0857   0.5411   0.1137
   1.750   0.6915   0.01226   0.00302  -0.0852   0.5309   0.1218
   2.000   0.7166   0.01230   0.00301  -0.0846   0.5207   0.1334
   2.500   0.7969   0.01102   0.00324  -0.0905   0.4964   1.0000
   2.750   0.8212   0.01120   0.00332  -0.0898   0.4846   1.0000
   3.000   0.8452   0.01138   0.00343  -0.0891   0.4718   1.0000
   3.250   0.8690   0.01158   0.00354  -0.0883   0.4575   1.0000
   3.500   0.8925   0.01178   0.00366  -0.0875   0.4422   1.0000
   3.750   0.9160   0.01199   0.00382  -0.0867   0.4270   1.0000
   4.000   0.9394   0.01222   0.00398  -0.0859   0.4125   1.0000
   4.250   0.9629   0.01245   0.00416  -0.0851   0.3999   1.0000
   4.500   0.9862   0.01270   0.00437  -0.0843   0.3877   1.0000
   4.750   1.0092   0.01297   0.00463  -0.0835   0.3759   1.0000
   5.000   1.0321   0.01325   0.00488  -0.0827   0.3654   1.0000
   5.250   1.0555   0.01353   0.00517  -0.0819   0.3560   1.0000
   5.500   1.0783   0.01383   0.00547  -0.0811   0.3473   1.0000
   5.750   1.1012   0.01414   0.00583  -0.0803   0.3380   1.0000
   6.000   1.1240   0.01445   0.00619  -0.0795   0.3290   1.0000
   6.500   1.1693   0.01511   0.00697  -0.0779   0.3121   1.0000
   6.750   1.1913   0.01548   0.00743  -0.0770   0.3035   1.0000
   7.000   1.2115   0.01587   0.00784  -0.0758   0.2813   1.0000
   7.250   1.2278   0.01649   0.00825  -0.0741   0.2345   1.0000
   7.500   1.2413   0.01743   0.00889  -0.0721   0.1811   1.0000
   7.750   1.2468   0.01917   0.01007  -0.0691   0.1045   1.0000
   8.000   1.2426   0.02166   0.01199  -0.0647   0.0179   1.0000
   8.250   1.2540   0.02280   0.01332  -0.0621   0.0125   1.0000
   8.500   1.2675   0.02367   0.01439  -0.0599   0.0108   1.0000
   8.750   1.2773   0.02469   0.01561  -0.0572   0.0097   1.0000
   9.000   1.2841   0.02585   0.01697  -0.0541   0.0090   1.0000
   9.250   1.2861   0.02732   0.01863  -0.0507   0.0082   1.0000
   9.500   1.2828   0.02914   0.02065  -0.0470   0.0076   1.0000
   9.750   1.2871   0.03053   0.02219  -0.0445   0.0069   1.0000
  10.000   1.2852   0.03244   0.02425  -0.0417   0.0066   1.0000
  10.250   1.2820   0.03462   0.02658  -0.0393   0.0064   1.0000
  10.500   1.2783   0.03702   0.02911  -0.0373   0.0063   1.0000
  10.750   1.2740   0.03970   0.03191  -0.0357   0.0060   1.0000
  11.000   1.2706   0.04249   0.03482  -0.0344   0.0059   1.0000
  11.250   1.2678   0.04539   0.03782  -0.0334   0.0058   1.0000
  11.500   1.2660   0.04832   0.04085  -0.0326   0.0057   1.0000
  11.750   1.2647   0.05129   0.04392  -0.0317   0.0055   1.0000
  12.000   1.2653   0.05417   0.04687  -0.0309   0.0054   1.0000
  12.250   1.2677   0.05688   0.04972  -0.0301   0.0053   1.0000
  12.500   1.2693   0.05989   0.05287  -0.0293   0.0048   1.0000
  12.750   1.2716   0.06286   0.05605  -0.0289   0.0046   1.0000
  13.000   1.2743   0.06587   0.05923  -0.0285   0.0045   1.0000
  13.250   1.2740   0.06924   0.06280  -0.0286   0.0043   1.0000
  13.500   1.2733   0.07294   0.06670  -0.0285   0.0042   1.0000
  13.750   1.2700   0.07706   0.07103  -0.0289   0.0041   1.0000
  14.000   1.2653   0.08136   0.07554  -0.0297   0.0041   1.0000
  14.250   1.2589   0.08610   0.08050  -0.0308   0.0041   1.0000
  14.500   1.2509   0.09115   0.08575  -0.0324   0.0041   1.0000
  14.750   1.2412   0.09665   0.09147  -0.0344   0.0041   1.0000
  15.000   1.2309   0.10242   0.09743  -0.0369   0.0041   1.0000
  15.250   1.2204   0.10840   0.10360  -0.0397   0.0041   1.0000
  15.500   1.2086   0.11487   0.11027  -0.0430   0.0041   1.0000
  15.750   1.1975   0.12141   0.11699  -0.0465   0.0042   1.0000
  16.000   1.1864   0.12821   0.12395  -0.0504   0.0042   1.0000
<< Back to GOE 380 AIRFOIL (goe380-il)

Polar data table (+)

Polar graphs


<< Back to GOE 380 AIRFOIL (goe380-il)