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

M6 (65%) (m665-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: M6 (65%) (m665-il)
Reynolds number: 100,000
Max Cl/Cd: 47.33 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m665-il-100000.txt
Download as CSV file: xf-m665-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: M6 (65%)                                        
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5374   0.08943   0.08456  -0.0104   1.0000   0.0958
  -8.000  -0.5471   0.08582   0.08101  -0.0134   1.0000   0.0983
  -7.750  -0.5701   0.08286   0.07799  -0.0190   1.0000   0.1009
  -7.500  -0.5674   0.07777   0.07293  -0.0191   1.0000   0.1028
  -7.250  -0.5504   0.07465   0.06989  -0.0162   1.0000   0.1085
  -7.000  -0.5641   0.07225   0.06718  -0.0201   1.0000   0.1152
  -6.750  -0.5480   0.06730   0.06246  -0.0176   1.0000   0.1202
  -6.500  -0.5563   0.06590   0.06058  -0.0192   1.0000   0.1300
  -6.250  -0.5374   0.06051   0.05551  -0.0172   1.0000   0.1343
  -6.000  -0.5349   0.05763   0.05237  -0.0170   1.0000   0.1458
  -5.750  -0.5209   0.05428   0.04913  -0.0153   1.0000   0.1521
  -5.500  -0.5132   0.05120   0.04594  -0.0141   1.0000   0.1636
  -5.250  -0.5044   0.04845   0.04309  -0.0124   1.0000   0.1775
  -5.000  -0.4943   0.04583   0.04042  -0.0105   1.0000   0.1927
  -4.750  -0.4866   0.04378   0.03828  -0.0081   1.0000   0.2194
  -4.500  -0.4582   0.03584   0.02886  -0.0070   1.0000   0.1136
  -4.250  -0.4398   0.03043   0.02264  -0.0036   1.0000   0.0838
  -4.000  -0.4224   0.02741   0.01918  -0.0011   1.0000   0.0790
  -3.750  -0.4037   0.02528   0.01667   0.0011   1.0000   0.0791
  -3.500  -0.3845   0.02374   0.01482   0.0031   1.0000   0.0824
  -3.250  -0.3638   0.02208   0.01286   0.0049   1.0000   0.0834
  -3.000  -0.3426   0.02074   0.01125   0.0067   1.0000   0.0855
  -2.750  -0.3226   0.01949   0.00997   0.0081   1.0000   0.0910
  -2.500  -0.3033   0.01862   0.00906   0.0097   1.0000   0.0972
  -2.250  -0.2843   0.01773   0.00811   0.0114   1.0000   0.1030
  -2.000  -0.0422   0.01249   0.00588  -0.0266   1.0000   1.0000
  -1.750  -0.0273   0.01244   0.00570  -0.0245   1.0000   1.0000
  -1.500  -0.0145   0.01245   0.00562  -0.0222   1.0000   1.0000
  -1.250  -0.0061   0.01257   0.00566  -0.0191   1.0000   1.0000
  -1.000  -0.0055   0.01284   0.00587  -0.0150   1.0000   1.0000
  -0.750  -0.0094   0.01325   0.00622  -0.0105   1.0000   1.0000
  -0.500   0.0098   0.01361   0.00649  -0.0104   0.9957   1.0000
  -0.250   0.0696   0.01375   0.00650  -0.0175   0.9829   1.0000
   0.000   0.1278   0.01379   0.00646  -0.0242   0.9701   1.0000
   0.250   0.1851   0.01374   0.00636  -0.0305   0.9570   1.0000
   0.500   0.2418   0.01360   0.00622  -0.0365   0.9434   1.0000
   0.750   0.2942   0.01344   0.00606  -0.0414   0.9283   1.0000
   1.000   0.3399   0.01330   0.00593  -0.0448   0.9108   1.0000
   1.250   0.3757   0.01325   0.00590  -0.0462   0.8902   1.0000
   1.500   0.4073   0.01323   0.00588  -0.0464   0.8704   1.0000
   1.750   0.4321   0.01333   0.00598  -0.0454   0.8494   1.0000
   2.000   0.4550   0.01345   0.00612  -0.0440   0.8291   1.0000
   2.250   0.4771   0.01357   0.00622  -0.0422   0.8094   1.0000
   2.500   0.4974   0.01375   0.00641  -0.0402   0.7878   1.0000
   2.750   0.5186   0.01389   0.00654  -0.0382   0.7678   1.0000
   3.000   0.5395   0.01408   0.00678  -0.0364   0.7471   1.0000
   3.250   0.5608   0.01426   0.00700  -0.0346   0.7271   1.0000
   3.500   0.5825   0.01443   0.00717  -0.0328   0.7077   1.0000
   3.750   0.6028   0.01456   0.00735  -0.0307   0.6828   1.0000
   4.000   0.6227   0.01459   0.00741  -0.0282   0.6542   1.0000
   4.250   0.6413   0.01451   0.00726  -0.0254   0.6184   1.0000
   4.500   0.6592   0.01444   0.00713  -0.0225   0.5750   1.0000
   4.750   0.6774   0.01448   0.00715  -0.0199   0.5265   1.0000
   5.000   0.6938   0.01466   0.00717  -0.0171   0.4546   1.0000
   5.250   0.6961   0.01649   0.00761  -0.0125   0.2028   1.0000
   5.500   0.7025   0.01904   0.00924  -0.0092   0.1041   1.0000
   5.750   0.7172   0.02044   0.01055  -0.0068   0.0874   1.0000
   6.000   0.7327   0.02199   0.01200  -0.0045   0.0797   1.0000
   6.250   0.7529   0.02336   0.01340  -0.0028   0.0744   1.0000
   6.500   0.7738   0.02515   0.01506  -0.0016   0.0695   1.0000
   6.750   0.7964   0.02686   0.01694  -0.0003   0.0662   1.0000
   7.000   0.8193   0.02865   0.01895   0.0010   0.0642   1.0000
   7.250   0.8415   0.03078   0.02136   0.0023   0.0633   1.0000
   7.500   0.8615   0.03322   0.02416   0.0040   0.0632   1.0000
   7.750   0.8790   0.03602   0.02736   0.0058   0.0639   1.0000
   8.000   0.8939   0.03921   0.03095   0.0078   0.0653   1.0000
   8.250   0.9058   0.04266   0.03480   0.0099   0.0663   1.0000
   8.500   0.9144   0.04604   0.03858   0.0122   0.0665   1.0000
   8.750   0.9212   0.05007   0.04291   0.0142   0.0673   1.0000
   9.000   0.9106   0.05519   0.04896   0.0183   0.0774   1.0000
<< Back to M6 (65%) (m665-il)

Polar data table (+)

Polar graphs


<< Back to M6 (65%) (m665-il)