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

RG 8 AIRFOIL (rg8-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: RG 8 AIRFOIL (rg8-il)
Reynolds number: 50,000
Max Cl/Cd: 35.51 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rg8-il-50000.txt
Download as CSV file: xf-rg8-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RG 8 AIRFOIL                                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4106   0.10765   0.10052  -0.0181   1.0000   0.2840
  -8.750  -0.4213   0.10571   0.09869  -0.0180   1.0000   0.2945
  -8.500  -0.4215   0.10329   0.09633  -0.0171   1.0000   0.3071
  -8.250  -0.4045   0.09879   0.09183  -0.0162   1.0000   0.3149
  -8.000  -0.4122   0.09642   0.08957  -0.0154   1.0000   0.3252
  -7.500  -0.4085   0.09029   0.08354  -0.0135   1.0000   0.3417
  -7.250  -0.5174   0.07680   0.07047  -0.0350   1.0000   0.1889
  -7.000  -0.5539   0.06248   0.05569  -0.0464   1.0000   0.1326
  -6.750  -0.5465   0.05820   0.05130  -0.0460   1.0000   0.1293
  -6.500  -0.5407   0.05366   0.04653  -0.0463   1.0000   0.1271
  -6.250  -0.5323   0.04893   0.04143  -0.0470   1.0000   0.1246
  -6.000  -0.5194   0.04415   0.03606  -0.0477   1.0000   0.1220
  -5.750  -0.5014   0.03997   0.03121  -0.0481   1.0000   0.1206
  -5.500  -0.4807   0.03677   0.02758  -0.0478   1.0000   0.1219
  -5.250  -0.4586   0.03414   0.02455  -0.0474   1.0000   0.1268
  -5.000  -0.4356   0.03174   0.02179  -0.0469   1.0000   0.1356
  -4.750  -0.4111   0.02960   0.01926  -0.0462   1.0000   0.1469
  -4.500  -0.3885   0.02773   0.01749  -0.0452   1.0000   0.1666
  -4.250  -0.3653   0.02599   0.01573  -0.0443   1.0000   0.2031
  -4.000  -0.3434   0.02430   0.01442  -0.0434   1.0000   0.2599
  -3.750  -0.3220   0.02257   0.01323  -0.0422   1.0000   0.3409
  -3.500  -0.3043   0.02142   0.01304  -0.0395   1.0000   0.4639
  -3.250  -0.2946   0.02160   0.01377  -0.0339   1.0000   0.5899
  -3.000  -0.2898   0.02199   0.01432  -0.0271   1.0000   0.6677
  -2.750  -0.2864   0.02220   0.01457  -0.0203   1.0000   0.7263
  -2.500  -0.2847   0.02223   0.01459  -0.0133   1.0000   0.7752
  -2.250  -0.2832   0.02207   0.01440  -0.0067   1.0000   0.8213
  -2.000  -0.2800   0.02178   0.01406  -0.0007   1.0000   0.8672
  -1.750  -0.2599   0.02153   0.01370   0.0022   1.0000   0.9214
  -1.500  -0.0954   0.02226   0.01357  -0.0218   1.0000   0.9987
  -1.250  -0.1044   0.02175   0.01303  -0.0186   1.0000   1.0000
  -1.000  -0.1156   0.02120   0.01246  -0.0150   1.0000   1.0000
  -0.750  -0.1123   0.02089   0.01205  -0.0137   1.0000   1.0000
  -0.500  -0.0948   0.02090   0.01189  -0.0146   1.0000   1.0000
  -0.250  -0.0721   0.02112   0.01192  -0.0162   1.0000   1.0000
   0.000  -0.0481   0.02147   0.01208  -0.0178   1.0000   1.0000
   0.250  -0.0243   0.02193   0.01237  -0.0193   1.0000   1.0000
   0.500  -0.0012   0.02247   0.01276  -0.0206   1.0000   1.0000
   0.750   0.0209   0.02309   0.01325  -0.0216   1.0000   1.0000
   1.000   0.0420   0.02378   0.01382  -0.0226   1.0000   1.0000
   1.250   0.0622   0.02454   0.01448  -0.0234   1.0000   1.0000
   1.500   0.1156   0.02609   0.01594  -0.0302   0.9846   1.0000
   1.750   0.1668   0.02733   0.01711  -0.0363   0.9654   1.0000
   2.000   0.2191   0.02847   0.01822  -0.0423   0.9469   1.0000
   2.250   0.2657   0.02937   0.01912  -0.0469   0.9274   1.0000
   2.500   0.3117   0.03015   0.01992  -0.0512   0.9077   1.0000
   2.750   0.3640   0.03083   0.02064  -0.0561   0.8889   1.0000
   3.000   0.4016   0.03138   0.02126  -0.0585   0.8676   1.0000
   3.250   0.4492   0.03175   0.02172  -0.0620   0.8476   1.0000
   3.500   0.4953   0.03197   0.02206  -0.0649   0.8275   1.0000
   3.750   0.5360   0.03213   0.02237  -0.0668   0.8059   1.0000
   4.000   0.5927   0.03164   0.02207  -0.0702   0.7871   1.0000
   4.250   0.6320   0.03151   0.02213  -0.0712   0.7646   1.0000
   4.500   0.6908   0.03048   0.02133  -0.0739   0.7442   1.0000
   4.750   0.7326   0.02997   0.02100  -0.0744   0.7204   1.0000
   5.000   0.7914   0.02853   0.01980  -0.0762   0.6973   1.0000
   5.250   0.8304   0.02796   0.01938  -0.0758   0.6693   1.0000
   5.500   0.8684   0.02747   0.01902  -0.0752   0.6393   1.0000
   5.750   0.8990   0.02738   0.01899  -0.0739   0.6070   1.0000
   6.000   0.9263   0.02748   0.01914  -0.0722   0.5726   1.0000
   6.250   0.9563   0.02750   0.01911  -0.0707   0.5371   1.0000
   6.500   0.9788   0.02794   0.01956  -0.0685   0.5001   1.0000
   6.750   1.0045   0.02829   0.01976  -0.0667   0.4620   1.0000
   7.000   1.0223   0.02906   0.02051  -0.0642   0.4231   1.0000
   7.250   1.0419   0.02986   0.02117  -0.0619   0.3840   1.0000
   7.500   1.0612   0.03081   0.02194  -0.0597   0.3448   1.0000
   7.750   1.0754   0.03208   0.02317  -0.0571   0.3069   1.0000
   8.000   1.0931   0.03342   0.02423  -0.0549   0.2705   1.0000
   8.250   1.1066   0.03514   0.02595  -0.0525   0.2379   1.0000
   8.500   1.1217   0.03709   0.02785  -0.0503   0.2085   1.0000
   8.750   1.1372   0.03921   0.02986  -0.0483   0.1822   1.0000
   9.000   1.1517   0.04140   0.03205  -0.0463   0.1590   1.0000
   9.250   1.1636   0.04413   0.03490  -0.0441   0.1406   1.0000
   9.500   1.1736   0.04699   0.03796  -0.0418   0.1258   1.0000
   9.750   1.1848   0.05005   0.04115  -0.0399   0.1139   1.0000
  10.000   1.1921   0.05318   0.04457  -0.0375   0.1054   1.0000
  10.250   1.1933   0.05706   0.04879  -0.0350   0.1004   1.0000
  10.500   1.1842   0.06101   0.05325  -0.0317   0.0980   1.0000
  10.750   1.1722   0.06517   0.05782  -0.0287   0.0969   1.0000
  11.000   1.1539   0.06922   0.06220  -0.0256   0.0966   1.0000
  11.250   1.1305   0.07349   0.06673  -0.0230   0.0971   1.0000
  11.500   1.1038   0.07828   0.07175  -0.0217   0.0979   1.0000
  11.750   1.0762   0.08378   0.07744  -0.0220   0.0991   1.0000
  12.000   1.0495   0.09011   0.08389  -0.0238   0.1003   1.0000
  12.250   1.0251   0.09724   0.09110  -0.0269   0.1015   1.0000
  12.500   1.0064   0.10477   0.09867  -0.0303   0.1025   1.0000
<< Back to RG 8 AIRFOIL (rg8-il)

Polar data table (+)

Polar graphs


<< Back to RG 8 AIRFOIL (rg8-il)