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

GOE 207 (AVIATIK V8) AIRFOIL (goe207-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: GOE 207 (AVIATIK V8) AIRFOIL (goe207-il)
Reynolds number: 1,000,000
Max Cl/Cd: 113.92 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe207-il-1000000-n5.txt
Download as CSV file: xf-goe207-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 207 (AVIATIK V8) 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.2511   0.09451   0.09140  -0.0111   0.5457   0.0135
  -7.500  -0.2425   0.09182   0.08867  -0.0127   0.5358   0.0140
  -7.250  -0.2342   0.08751   0.08434  -0.0172   0.5299   0.0149
  -7.000  -0.2201   0.08511   0.08190  -0.0194   0.5212   0.0150
  -6.750  -0.2045   0.08264   0.07940  -0.0220   0.5139   0.0152
  -6.500  -0.1875   0.08001   0.07675  -0.0251   0.5061   0.0155
  -6.250  -0.1683   0.07744   0.07413  -0.0284   0.4979   0.0169
  -6.000  -0.1476   0.07400   0.07064  -0.0329   0.4903   0.0171
  -5.750  -0.1255   0.07059   0.06717  -0.0374   0.4824   0.0171
  -5.500  -0.0996   0.06683   0.06335  -0.0429   0.4757   0.0173
  -5.250  -0.0733   0.06330   0.05974  -0.0478   0.4684   0.0173
  -5.000  -0.0442   0.05973   0.05608  -0.0529   0.4625   0.0173
  -4.750  -0.0138   0.05623   0.05249  -0.0578   0.4554   0.0174
  -4.500   0.0079   0.05362   0.04983  -0.0595   0.4476   0.0175
  -4.250   0.0328   0.05144   0.04759  -0.0617   0.4400   0.0177
  -4.000   0.0597   0.04934   0.04541  -0.0641   0.4320   0.0179
  -3.750   0.0878   0.04732   0.04332  -0.0666   0.4250   0.0183
  -3.500   0.1214   0.04493   0.04081  -0.0698   0.4167   0.0200
  -3.250   0.1560   0.04208   0.03783  -0.0731   0.4094   0.0201
  -3.000   0.1901   0.03941   0.03501  -0.0761   0.4008   0.0202
  -2.750   0.2230   0.03698   0.03244  -0.0784   0.3929   0.0202
  -2.500   0.2547   0.03476   0.03007  -0.0804   0.3838   0.0202
  -2.250   0.2873   0.03257   0.02774  -0.0822   0.3759   0.0202
  -1.750   0.3459   0.02914   0.02407  -0.0849   0.3576   0.0199
  -1.500   0.3772   0.02745   0.02223  -0.0862   0.3488   0.0197
  -1.250   0.4107   0.02549   0.02009  -0.0875   0.3407   0.0199
  -0.500   0.5099   0.02014   0.01416  -0.0903   0.3210   0.0205
  -0.250   0.5385   0.01972   0.01367  -0.0909   0.3155   0.0208
   0.000   0.5742   0.01697   0.01058  -0.0917   0.3105   0.0205
   0.250   0.6049   0.01589   0.00931  -0.0922   0.3048   0.0207
   0.500   0.6348   0.01517   0.00847  -0.0925   0.2998   0.0209
   0.750   0.6648   0.01441   0.00756  -0.0928   0.2934   0.0213
   1.000   0.6955   0.01308   0.00598  -0.0931   0.2863   0.0219
   1.250   0.7253   0.01217   0.00487  -0.0934   0.2817   0.0223
   1.500   0.7544   0.01178   0.00436  -0.0935   0.2767   0.0227
   1.750   0.7831   0.01159   0.00407  -0.0937   0.2718   0.0232
   2.000   0.8120   0.01141   0.00384  -0.0938   0.2683   0.0234
   2.250   0.8408   0.01131   0.00370  -0.0940   0.2649   0.0236
   2.500   0.8696   0.01113   0.00348  -0.0942   0.2610   0.0241
   2.750   0.8982   0.01113   0.00345  -0.0944   0.2561   0.0246
   3.000   0.9268   0.01113   0.00345  -0.0946   0.2521   0.0251
   3.250   0.9554   0.01114   0.00345  -0.0947   0.2489   0.0256
   3.500   0.9839   0.01117   0.00346  -0.0949   0.2457   0.0262
   3.750   1.0123   0.01122   0.00350  -0.0950   0.2424   0.0268
   4.000   1.0405   0.01132   0.00359  -0.0951   0.2389   0.0275
   4.250   1.0689   0.01134   0.00363  -0.0953   0.2363   0.0287
   4.500   1.0971   0.01143   0.00371  -0.0954   0.2320   0.0297
   4.750   1.1250   0.01155   0.00381  -0.0955   0.2271   0.0309
   5.000   1.1528   0.01171   0.00395  -0.0956   0.2222   0.0322
   5.250   1.1807   0.01181   0.00406  -0.0957   0.2187   0.0358
   5.500   1.2084   0.01194   0.00424  -0.0958   0.2137   0.0664
   5.750   1.2358   0.01213   0.00441  -0.0959   0.2087   0.0791
   6.000   1.2632   0.01228   0.00459  -0.0960   0.2042   0.0943
   6.500   1.3123   0.01152   0.00520  -0.0954   0.1849   1.0000
   6.750   1.3319   0.01291   0.00614  -0.0947   0.1176   1.0000
   7.000   1.3561   0.01352   0.00666  -0.0944   0.1014   1.0000
   7.250   1.3806   0.01405   0.00712  -0.0942   0.0891   1.0000
   7.500   1.3983   0.01552   0.00831  -0.0931   0.0287   1.0000
   7.750   1.4230   0.01594   0.00872  -0.0929   0.0253   1.0000
   8.000   1.4477   0.01632   0.00912  -0.0926   0.0230   1.0000
   8.250   1.4720   0.01673   0.00953  -0.0923   0.0211   1.0000
   8.500   1.4956   0.01719   0.01000  -0.0919   0.0194   1.0000
   8.750   1.5194   0.01760   0.01043  -0.0915   0.0185   1.0000
   9.000   1.5427   0.01804   0.01089  -0.0911   0.0174   1.0000
   9.250   1.5652   0.01853   0.01138  -0.0906   0.0163   1.0000
   9.500   1.5869   0.01906   0.01193  -0.0900   0.0153   1.0000
   9.750   1.6087   0.01955   0.01245  -0.0894   0.0147   1.0000
  10.000   1.6297   0.02008   0.01301  -0.0887   0.0140   1.0000
  10.250   1.6497   0.02065   0.01361  -0.0879   0.0134   1.0000
  10.500   1.6685   0.02129   0.01427  -0.0869   0.0128   1.0000
  10.750   1.6852   0.02204   0.01506  -0.0857   0.0121   1.0000
  11.000   1.7018   0.02269   0.01576  -0.0844   0.0118   1.0000
  11.250   1.7165   0.02341   0.01653  -0.0829   0.0114   1.0000
  11.500   1.7277   0.02421   0.01737  -0.0809   0.0111   1.0000
  11.750   1.7352   0.02517   0.01840  -0.0785   0.0107   1.0000
  12.000   1.7422   0.02634   0.01963  -0.0765   0.0104   1.0000
  12.250   1.7485   0.02776   0.02111  -0.0749   0.0101   1.0000
  12.500   1.7534   0.02947   0.02290  -0.0737   0.0098   1.0000
  12.750   1.7563   0.03160   0.02511  -0.0728   0.0096   1.0000
  13.000   1.7566   0.03423   0.02784  -0.0723   0.0093   1.0000
  13.250   1.7584   0.03689   0.03060  -0.0721   0.0092   1.0000
  13.500   1.7577   0.04000   0.03381  -0.0723   0.0091   1.0000
  13.750   1.7544   0.04356   0.03748  -0.0726   0.0090   1.0000
  14.000   1.7481   0.04761   0.04165  -0.0732   0.0089   1.0000
  14.250   1.7387   0.05218   0.04634  -0.0739   0.0088   1.0000
  14.500   1.7261   0.05724   0.05153  -0.0748   0.0087   1.0000
  14.750   1.7102   0.06273   0.05715  -0.0757   0.0086   1.0000
  15.000   1.6918   0.06858   0.06313  -0.0767   0.0086   1.0000
  15.250   1.6711   0.07482   0.06950  -0.0778   0.0085   1.0000
  15.500   1.6513   0.08111   0.07592  -0.0791   0.0085   1.0000
  15.750   1.6304   0.08773   0.08266  -0.0805   0.0084   1.0000
<< Back to GOE 207 (AVIATIK V8) AIRFOIL (goe207-il)

Polar data table (+)

Polar graphs


<< Back to GOE 207 (AVIATIK V8) AIRFOIL (goe207-il)