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

GOE 376 AIRFOIL (goe376-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 376 AIRFOIL (goe376-il)
Reynolds number: 200,000
Max Cl/Cd: 76.5 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe376-il-200000-n5.txt
Download as CSV file: xf-goe376-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 376 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.2326   0.09285   0.08962  -0.0269   1.0000   0.0194
  -8.500  -0.2329   0.09011   0.08691  -0.0271   1.0000   0.0195
  -8.250  -0.2341   0.08747   0.08432  -0.0273   1.0000   0.0195
  -8.000  -0.2397   0.08532   0.08223  -0.0275   1.0000   0.0195
  -7.750  -0.2315   0.08072   0.07766  -0.0255   1.0000   0.0199
  -7.500  -0.2315   0.07804   0.07503  -0.0243   1.0000   0.0203
  -7.250  -0.2358   0.07585   0.07291  -0.0228   1.0000   0.0205
  -7.000  -0.2271   0.07219   0.06926  -0.0252   0.9959   0.0209
  -6.750  -0.2133   0.06802   0.06509  -0.0294   0.9878   0.0214
  -6.500  -0.2950   0.08151   0.07843  -0.0242   0.9967   0.0203
  -6.250  -0.2684   0.07728   0.07418  -0.0303   0.9901   0.0209
  -6.000  -0.2398   0.07302   0.06988  -0.0370   0.9818   0.0218
  -5.750  -0.2065   0.06868   0.06548  -0.0453   0.9711   0.0236
  -5.500  -0.1571   0.06396   0.06056  -0.0586   0.9578   0.0244
  -5.250  -0.1218   0.05939   0.05583  -0.0649   0.9471   0.0245
  -5.000  -0.1006   0.05461   0.05105  -0.0673   0.9387   0.0249
  -4.750  -0.0748   0.05119   0.04759  -0.0700   0.9290   0.0256
  -4.500  -0.0397   0.04787   0.04416  -0.0748   0.9211   0.0279
  -4.250   0.0176   0.04461   0.04043  -0.0828   0.9115   0.0309
  -4.000   0.0510   0.04029   0.03591  -0.0864   0.9016   0.0313
  -3.750   0.0765   0.03689   0.03249  -0.0888   0.8914   0.0324
  -3.500   0.1059   0.03469   0.03018  -0.0908   0.8788   0.0348
  -3.250   0.1466   0.03335   0.02834  -0.0926   0.8653   0.0397
  -3.000   0.1729   0.03007   0.02480  -0.0935   0.8513   0.0407
  -2.750   0.1958   0.02781   0.02248  -0.0939   0.8366   0.0420
  -2.500   0.2205   0.02637   0.02091  -0.0940   0.8212   0.0451
  -2.250   0.2494   0.02494   0.01905  -0.0938   0.8060   0.0528
  -2.000   0.2729   0.02345   0.01747  -0.0936   0.7907   0.0551
  -1.750   0.2987   0.02245   0.01625  -0.0932   0.7759   0.0602
  -0.750   0.4049   0.01780   0.01025  -0.0898   0.7184   0.0524
  -0.500   0.4302   0.01655   0.00888  -0.0891   0.7041   0.0452
  -0.250   0.4575   0.01624   0.00818  -0.0880   0.6891   0.0403
   0.000   0.4825   0.01536   0.00715  -0.0873   0.6718   0.0388
   0.250   0.5073   0.01476   0.00636  -0.0864   0.6509   0.0382
   0.500   0.5319   0.01430   0.00572  -0.0854   0.6276   0.0378
   0.750   0.5565   0.01393   0.00519  -0.0845   0.6056   0.0377
   1.000   0.5812   0.01362   0.00476  -0.0836   0.5860   0.0379
   1.250   0.6058   0.01349   0.00451  -0.0828   0.5665   0.0392
   1.500   0.6299   0.01333   0.00425  -0.0819   0.5473   0.0404
   1.750   0.6540   0.01316   0.00401  -0.0810   0.5284   0.0403
   2.000   0.6779   0.01305   0.00385  -0.0802   0.5115   0.0402
   2.250   0.7018   0.01297   0.00372  -0.0793   0.4957   0.0402
   2.500   0.7256   0.01294   0.00363  -0.0784   0.4811   0.0403
   2.750   0.7496   0.01296   0.00358  -0.0776   0.4674   0.0406
   3.000   0.7738   0.01302   0.00356  -0.0768   0.4553   0.0412
   3.500   0.8227   0.01325   0.00366  -0.0754   0.4315   0.0436
   3.750   0.8472   0.01340   0.00377  -0.0747   0.4194   0.0461
   4.000   0.8714   0.01355   0.00391  -0.0740   0.4064   0.0554
   4.500   0.9534   0.01263   0.00452  -0.0805   0.3738   1.0000
   4.750   0.9772   0.01288   0.00473  -0.0797   0.3615   1.0000
   5.000   1.0008   0.01314   0.00495  -0.0789   0.3491   1.0000
   5.250   1.0242   0.01341   0.00521  -0.0781   0.3359   1.0000
   5.500   1.0473   0.01369   0.00547  -0.0772   0.3218   1.0000
   5.750   1.0701   0.01400   0.00576  -0.0763   0.3068   1.0000
   6.000   1.0925   0.01434   0.00607  -0.0754   0.2923   1.0000
   6.250   1.1144   0.01472   0.00642  -0.0744   0.2755   1.0000
   6.500   1.1353   0.01516   0.00680  -0.0732   0.2561   1.0000
   6.750   1.1554   0.01566   0.00721  -0.0720   0.2365   1.0000
   7.000   1.1754   0.01619   0.00765  -0.0708   0.2166   1.0000
   7.250   1.1957   0.01669   0.00813  -0.0696   0.1984   1.0000
   7.500   1.2158   0.01720   0.00861  -0.0685   0.1800   1.0000
   7.750   1.2341   0.01785   0.00913  -0.0671   0.1521   1.0000
   8.000   1.2515   0.01859   0.00975  -0.0656   0.1256   1.0000
   8.250   1.2656   0.01963   0.01059  -0.0637   0.0865   1.0000
   8.500   1.2662   0.02188   0.01229  -0.0600   0.0220   1.0000
   8.750   1.2800   0.02287   0.01334  -0.0579   0.0184   1.0000
   9.000   1.2939   0.02380   0.01441  -0.0558   0.0165   1.0000
   9.250   1.3062   0.02481   0.01561  -0.0535   0.0152   1.0000
   9.500   1.3171   0.02578   0.01674  -0.0510   0.0145   1.0000
   9.750   1.3260   0.02676   0.01791  -0.0483   0.0136   1.0000
  10.000   1.3330   0.02786   0.01917  -0.0455   0.0128   1.0000
  10.250   1.3381   0.02910   0.02057  -0.0427   0.0123   1.0000
  10.500   1.3404   0.03058   0.02220  -0.0398   0.0118   1.0000
  10.750   1.3405   0.03229   0.02406  -0.0371   0.0114   1.0000
  11.000   1.3390   0.03423   0.02615  -0.0347   0.0111   1.0000
  11.250   1.3355   0.03651   0.02857  -0.0327   0.0109   1.0000
  11.500   1.3309   0.03909   0.03128  -0.0311   0.0107   1.0000
  11.750   1.3238   0.04216   0.03448  -0.0299   0.0105   1.0000
  12.000   1.3180   0.04532   0.03777  -0.0292   0.0104   1.0000
  12.250   1.3101   0.04892   0.04149  -0.0287   0.0102   1.0000
  12.500   1.3025   0.05262   0.04529  -0.0283   0.0100   1.0000
  12.750   1.2994   0.05590   0.04866  -0.0281   0.0099   1.0000
  13.000   1.2988   0.05899   0.05188  -0.0281   0.0098   1.0000
  13.250   1.2982   0.06217   0.05519  -0.0284   0.0096   1.0000
  13.500   1.2968   0.06557   0.05876  -0.0289   0.0093   1.0000
  13.750   1.2950   0.06910   0.06244  -0.0296   0.0090   1.0000
  14.000   1.2924   0.07277   0.06624  -0.0301   0.0087   1.0000
  14.250   1.2897   0.07652   0.07012  -0.0308   0.0085   1.0000
  14.500   1.2868   0.08040   0.07413  -0.0317   0.0084   1.0000
  14.750   1.2838   0.08438   0.07825  -0.0325   0.0083   1.0000
  15.000   1.2803   0.08851   0.08252  -0.0336   0.0081   1.0000
  15.250   1.2761   0.09286   0.08703  -0.0349   0.0080   1.0000
  15.500   1.2714   0.09741   0.09173  -0.0364   0.0079   1.0000
  15.750   1.2658   0.10223   0.09670  -0.0382   0.0078   1.0000
  16.000   1.2596   0.10726   0.10188  -0.0402   0.0078   1.0000
  16.250   1.2526   0.11257   0.10734  -0.0425   0.0077   1.0000
  16.500   1.2450   0.11816   0.11310  -0.0452   0.0077   1.0000
  16.750   1.2367   0.12404   0.11914  -0.0482   0.0077   1.0000
  17.000   1.2277   0.13023   0.12550  -0.0515   0.0076   1.0000
  17.250   1.2182   0.13673   0.13215  -0.0551   0.0076   1.0000
  17.500   1.2079   0.14369   0.13928  -0.0592   0.0077   1.0000
  17.750   1.1975   0.15090   0.14665  -0.0635   0.0077   1.0000
  18.000   1.1863   0.15865   0.15455  -0.0683   0.0077   1.0000
  18.250   1.1741   0.16696   0.16302  -0.0736   0.0078   1.0000
  18.500   1.1612   0.17589   0.17210  -0.0792   0.0079   1.0000
  18.750   1.1468   0.18584   0.18218  -0.0854   0.0080   1.0000
  19.000   1.1309   0.19714   0.19358  -0.0922   0.0082   1.0000
<< Back to GOE 376 AIRFOIL (goe376-il)

Polar data table (+)

Polar graphs


<< Back to GOE 376 AIRFOIL (goe376-il)