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

GIII BL126 AIRFOIL (giiie-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GIII BL126 AIRFOIL (giiie-il)
Reynolds number: 500,000
Max Cl/Cd: 76.75 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-giiie-il-500000.txt
Download as CSV file: xf-giiie-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GIII BL126 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4816   0.09550   0.09344  -0.0068   1.0000   0.0177
  -9.750  -0.6088   0.09724   0.09510  -0.0010   1.0000   0.0157
  -9.500  -0.6150   0.09159   0.08948  -0.0041   1.0000   0.0161
  -9.250  -0.6159   0.08726   0.08517  -0.0066   1.0000   0.0163
  -9.000  -0.6192   0.08182   0.07976  -0.0113   1.0000   0.0164
  -8.750  -0.6288   0.07527   0.07319  -0.0181   1.0000   0.0164
  -8.500  -0.6335   0.07071   0.06858  -0.0204   1.0000   0.0166
  -8.250  -0.6330   0.06649   0.06430  -0.0222   1.0000   0.0169
  -8.000  -0.6297   0.06254   0.06027  -0.0234   1.0000   0.0173
  -7.750  -0.6245   0.05848   0.05611  -0.0242   1.0000   0.0179
  -7.500  -0.6176   0.05417   0.05167  -0.0248   1.0000   0.0188
  -7.250  -0.6088   0.04950   0.04681  -0.0248   1.0000   0.0200
  -7.000  -0.5864   0.04617   0.04314  -0.0240   1.0000   0.0225
  -6.750  -0.5770   0.04222   0.03889  -0.0227   1.0000   0.0226
  -6.500  -0.5867   0.03324   0.02942  -0.0201   1.0000   0.0239
  -6.250  -0.5832   0.02520   0.02055  -0.0150   1.0000   0.0157
  -6.000  -0.5651   0.02427   0.01948  -0.0132   0.9998   0.0154
  -5.750  -0.5357   0.01907   0.01365  -0.0143   0.9942   0.0151
  -5.500  -0.5029   0.01647   0.01066  -0.0155   0.9874   0.0158
  -5.250  -0.4704   0.01416   0.00813  -0.0167   0.9799   0.0169
  -5.000  -0.4358   0.01349   0.00743  -0.0185   0.9706   0.0190
  -4.750  -0.4034   0.01267   0.00652  -0.0195   0.9588   0.0208
  -4.500  -0.3732   0.01199   0.00571  -0.0200   0.9452   0.0222
  -4.250  -0.3480   0.01100   0.00460  -0.0194   0.9301   0.0244
  -4.000  -0.3230   0.01052   0.00407  -0.0188   0.9164   0.0286
  -3.750  -0.2975   0.01019   0.00364  -0.0182   0.9044   0.0327
  -3.500  -0.2735   0.00959   0.00306  -0.0174   0.8937   0.0495
  -3.000  -0.2227   0.00901   0.00259  -0.0165   0.8752   0.1088
  -2.750  -0.1968   0.00877   0.00241  -0.0162   0.8676   0.1376
  -2.500  -0.1717   0.00841   0.00224  -0.0159   0.8599   0.1941
  -2.250  -0.1500   0.00756   0.00198  -0.0152   0.8528   0.3618
  -2.000  -0.1368   0.00620   0.00184  -0.0125   0.8451   0.7003
  -1.750  -0.1130   0.00606   0.00187  -0.0114   0.8389   0.7679
  -1.500  -0.0877   0.00602   0.00187  -0.0106   0.8324   0.8029
  -1.250  -0.0620   0.00602   0.00186  -0.0099   0.8266   0.8279
  -1.000  -0.0357   0.00601   0.00186  -0.0094   0.8201   0.8455
  -0.750  -0.0099   0.00604   0.00187  -0.0087   0.8141   0.8633
  -0.500   0.0162   0.00604   0.00188  -0.0081   0.8059   0.8781
  -0.250   0.0419   0.00608   0.00188  -0.0074   0.7980   0.8926
   0.000   0.0676   0.00609   0.00190  -0.0067   0.7896   0.9067
   0.250   0.0932   0.00614   0.00194  -0.0059   0.7817   0.9208
   0.500   0.1176   0.00622   0.00200  -0.0047   0.7734   0.9388
   0.750   0.1459   0.00628   0.00207  -0.0044   0.7648   0.9503
   1.000   0.1753   0.00633   0.00206  -0.0047   0.7578   0.9568
   1.250   0.2080   0.00634   0.00209  -0.0057   0.7494   0.9615
   1.500   0.2382   0.00638   0.00211  -0.0062   0.7419   0.9680
   1.750   0.2729   0.00639   0.00213  -0.0078   0.7331   0.9716
   2.000   0.3086   0.00641   0.00215  -0.0095   0.7238   0.9750
   2.250   0.3433   0.00643   0.00217  -0.0111   0.7140   0.9791
   2.500   0.3773   0.00645   0.00219  -0.0125   0.7016   0.9834
   2.750   0.4140   0.00643   0.00217  -0.0145   0.6796   0.9862
   3.000   0.4493   0.00648   0.00216  -0.0162   0.6517   0.9900
   3.250   0.4833   0.00657   0.00221  -0.0176   0.6230   0.9942
   3.500   0.5188   0.00676   0.00225  -0.0195   0.5762   0.9970
   3.750   0.5519   0.00720   0.00234  -0.0211   0.4782   1.0000
   4.000   0.5720   0.00796   0.00259  -0.0202   0.3612   1.0000
   4.250   0.5928   0.00851   0.00287  -0.0193   0.2912   1.0000
   4.500   0.6128   0.00910   0.00315  -0.0183   0.2176   1.0000
   4.750   0.6304   0.00997   0.00355  -0.0169   0.1203   1.0000
   5.000   0.6479   0.01083   0.00405  -0.0154   0.0553   1.0000
   5.250   0.6678   0.01144   0.00460  -0.0140   0.0358   1.0000
   5.500   0.6879   0.01207   0.00527  -0.0126   0.0289   1.0000
   5.750   0.7095   0.01250   0.00573  -0.0116   0.0251   1.0000
   6.000   0.7279   0.01336   0.00665  -0.0099   0.0219   1.0000
   6.250   0.7463   0.01431   0.00769  -0.0083   0.0204   1.0000
   6.500   0.7678   0.01490   0.00835  -0.0072   0.0189   1.0000
   6.750   0.7898   0.01546   0.00896  -0.0063   0.0170   1.0000
   7.000   0.8105   0.01628   0.00984  -0.0052   0.0157   1.0000
   7.250   0.8277   0.01797   0.01161  -0.0036   0.0146   1.0000
   7.500   0.8460   0.02022   0.01406  -0.0021   0.0139   1.0000
   7.750   0.8679   0.02143   0.01542  -0.0012   0.0134   1.0000
   8.000   0.8890   0.02296   0.01715  -0.0001   0.0128   1.0000
   8.250   0.9097   0.02426   0.01863   0.0008   0.0120   1.0000
   8.500   0.9293   0.02575   0.02030   0.0018   0.0113   1.0000
   8.750   0.9461   0.02802   0.02286   0.0032   0.0110   1.0000
   9.000   0.9598   0.03075   0.02591   0.0048   0.0108   1.0000
   9.250   0.9630   0.03576   0.03149   0.0075   0.0111   1.0000
   9.500   0.9467   0.04381   0.04021   0.0114   0.0125   1.0000
<< Back to GIII BL126 AIRFOIL (giiie-il)

Polar data table (+)

Polar graphs


<< Back to GIII BL126 AIRFOIL (giiie-il)