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

GOE 360 AIRFOIL (goe360-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 360 AIRFOIL (goe360-il)
Reynolds number: 1,000,000
Max Cl/Cd: 131.22 at α=7°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe360-il-1000000.txt
Download as CSV file: xf-goe360-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 360 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.1666   0.09083   0.08901  -0.0370   0.8688   0.0150
  -9.000  -0.1649   0.08736   0.08551  -0.0385   0.8623   0.0151
  -8.750  -0.1647   0.08366   0.08181  -0.0404   0.8549   0.0151
  -8.500  -0.1610   0.07954   0.07766  -0.0407   0.8474   0.0153
  -8.250  -0.1535   0.07646   0.07456  -0.0407   0.8382   0.0154
  -8.000  -0.1472   0.07362   0.07170  -0.0412   0.8288   0.0156
  -7.750  -0.2283   0.08368   0.08174  -0.0390   0.8503   0.0153
  -7.500  -0.2184   0.08099   0.07904  -0.0403   0.8410   0.0155
  -7.250  -0.2057   0.07822   0.07623  -0.0428   0.8311   0.0156
  -7.000  -0.1912   0.07532   0.07329  -0.0461   0.8197   0.0159
  -6.750  -0.1749   0.07226   0.07018  -0.0498   0.8074   0.0163
  -6.500  -0.1568   0.06901   0.06687  -0.0538   0.7936   0.0170
  -6.250  -0.1262   0.06410   0.06184  -0.0635   0.7778   0.0183
  -6.000  -0.0971   0.05915   0.05674  -0.0711   0.7601   0.0184
  -5.750  -0.0728   0.05526   0.05269  -0.0750   0.7398   0.0185
  -5.500  -0.0598   0.05177   0.04911  -0.0759   0.7175   0.0188
  -5.000  -0.0157   0.04713   0.04422  -0.0795   0.6788   0.0193
  -4.750   0.0095   0.04470   0.04166  -0.0817   0.6644   0.0198
  -4.500   0.0472   0.04144   0.03817  -0.0857   0.6531   0.0221
  -4.250   0.0787   0.03820   0.03472  -0.0880   0.6440   0.0222
  -4.000   0.1066   0.03540   0.03174  -0.0894   0.6356   0.0222
  -3.750   0.1301   0.03120   0.02737  -0.0911   0.6286   0.0226
  -3.500   0.1542   0.02971   0.02579  -0.0918   0.6206   0.0229
  -3.250   0.1799   0.02825   0.02425  -0.0924   0.6135   0.0233
  -3.000   0.2069   0.02671   0.02258  -0.0931   0.6065   0.0239
  -2.750   0.2350   0.02509   0.02081  -0.0936   0.6003   0.0250
  -2.500   0.2688   0.02375   0.01919  -0.0933   0.5940   0.0267
  -2.000   0.3240   0.01888   0.01384  -0.0940   0.5812   0.0273
  -1.750   0.3508   0.01785   0.01271  -0.0943   0.5738   0.0277
  -1.500   0.3783   0.01701   0.01179  -0.0945   0.5671   0.0282
  -1.250   0.4060   0.01622   0.01089  -0.0947   0.5594   0.0290
  -1.000   0.4345   0.01546   0.00998  -0.0947   0.5517   0.0307
  -0.750   0.4643   0.01565   0.00996  -0.0943   0.5423   0.0324
  -0.500   0.4927   0.01497   0.00911  -0.0943   0.5331   0.0325
   0.000   0.5477   0.01240   0.00630  -0.0947   0.5128   0.0346
   0.250   0.5756   0.01203   0.00585  -0.0948   0.5013   0.0359
   0.500   0.6036   0.01177   0.00546  -0.0947   0.4886   0.0379
   0.750   0.6318   0.01229   0.00584  -0.0945   0.4749   0.0396
   1.000   0.6590   0.01088   0.00433  -0.0946   0.4624   0.0421
   1.250   0.6865   0.01067   0.00406  -0.0946   0.4503   0.0442
   1.500   0.7144   0.01022   0.00351  -0.0944   0.4396   0.0422
   1.750   0.7416   0.01025   0.00354  -0.0946   0.4290   0.0522
   2.250   0.7969   0.00939   0.00249  -0.0941   0.4098   0.0342
   2.500   0.8243   0.00936   0.00242  -0.0941   0.4011   0.0350
   2.750   0.8520   0.00935   0.00239  -0.0941   0.3931   0.0361
   3.000   0.8795   0.00938   0.00239  -0.0941   0.3857   0.0373
   3.250   0.9071   0.00941   0.00240  -0.0942   0.3787   0.0385
   3.500   0.9344   0.00950   0.00246  -0.0942   0.3717   0.0395
   3.750   0.9621   0.00951   0.00247  -0.0942   0.3654   0.0435
   4.000   0.9892   0.00962   0.00255  -0.0942   0.3588   0.0481
   4.250   1.0128   0.00807   0.00285  -0.0939   0.3541   1.0000
   4.500   1.0401   0.00822   0.00295  -0.0939   0.3489   1.0000
   4.750   1.0669   0.00841   0.00309  -0.0938   0.3433   1.0000
   5.000   1.0943   0.00853   0.00322  -0.0939   0.3392   1.0000
   5.250   1.1214   0.00868   0.00335  -0.0939   0.3345   1.0000
   5.750   1.1749   0.00903   0.00365  -0.0938   0.3219   1.0000
   6.000   1.2013   0.00923   0.00382  -0.0938   0.3151   1.0000
   6.250   1.2279   0.00940   0.00399  -0.0937   0.3094   1.0000
   6.500   1.2545   0.00957   0.00415  -0.0937   0.3033   1.0000
   6.750   1.2804   0.00980   0.00435  -0.0936   0.2966   1.0000
   7.000   1.3069   0.00996   0.00453  -0.0936   0.2894   1.0000
   7.250   1.3325   0.01021   0.00475  -0.0934   0.2818   1.0000
   7.500   1.3584   0.01041   0.00494  -0.0933   0.2729   1.0000
   7.750   1.3837   0.01067   0.00518  -0.0931   0.2622   1.0000
   8.000   1.4077   0.01104   0.00547  -0.0928   0.2443   1.0000
   8.250   1.4273   0.01180   0.00598  -0.0919   0.2037   1.0000
   8.500   1.4235   0.01469   0.00799  -0.0878   0.0757   1.0000
   8.750   1.4301   0.01646   0.00943  -0.0850   0.0216   1.0000
   9.000   1.4497   0.01704   0.01003  -0.0840   0.0188   1.0000
   9.250   1.4671   0.01774   0.01078  -0.0827   0.0165   1.0000
   9.500   1.4858   0.01829   0.01137  -0.0816   0.0156   1.0000
   9.750   1.5024   0.01890   0.01203  -0.0802   0.0147   1.0000
  10.000   1.5155   0.01959   0.01276  -0.0782   0.0140   1.0000
  10.250   1.5257   0.02049   0.01373  -0.0759   0.0132   1.0000
  10.500   1.5312   0.02175   0.01509  -0.0733   0.0125   1.0000
  10.750   1.5432   0.02264   0.01604  -0.0718   0.0123   1.0000
  11.000   1.5538   0.02369   0.01715  -0.0702   0.0119   1.0000
  11.250   1.5633   0.02486   0.01840  -0.0687   0.0115   1.0000
  11.500   1.5718   0.02619   0.01978  -0.0673   0.0111   1.0000
  11.750   1.5789   0.02767   0.02133  -0.0659   0.0107   1.0000
  12.000   1.5838   0.02942   0.02316  -0.0646   0.0104   1.0000
  12.250   1.5843   0.03168   0.02551  -0.0632   0.0101   1.0000
  12.500   1.5780   0.03477   0.02872  -0.0619   0.0098   1.0000
  12.750   1.5644   0.03890   0.03300  -0.0609   0.0096   1.0000
  13.000   1.5670   0.04145   0.03563  -0.0606   0.0095   1.0000
  13.250   1.5675   0.04430   0.03858  -0.0604   0.0093   1.0000
  13.500   1.5658   0.04749   0.04187  -0.0603   0.0092   1.0000
  13.750   1.5620   0.05098   0.04546  -0.0603   0.0091   1.0000
  14.000   1.5559   0.05481   0.04939  -0.0604   0.0089   1.0000
  14.250   1.5494   0.05879   0.05347  -0.0606   0.0088   1.0000
  14.500   1.5419   0.06299   0.05777  -0.0610   0.0087   1.0000
  14.750   1.5339   0.06731   0.06218  -0.0614   0.0086   1.0000
  15.000   1.5255   0.07176   0.06673  -0.0620   0.0084   1.0000
  15.250   1.5176   0.07622   0.07128  -0.0626   0.0083   1.0000
  15.500   1.5091   0.08082   0.07597  -0.0634   0.0082   1.0000
  15.750   1.5018   0.08530   0.08054  -0.0642   0.0081   1.0000
<< Back to GOE 360 AIRFOIL (goe360-il)

Polar data table (+)

Polar graphs


<< Back to GOE 360 AIRFOIL (goe360-il)