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

RUTAN CANARD AIRFOIL (amsoil1-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: RUTAN CANARD AIRFOIL (amsoil1-il)
Reynolds number: 50,000
Max Cl/Cd: 33.02 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-amsoil1-il-50000.txt
Download as CSV file: xf-amsoil1-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RUTAN CANARD 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.4602   0.12897   0.12133  -0.0120   1.0000   0.2218
 -10.250  -0.4441   0.12517   0.11751  -0.0116   1.0000   0.2324
 -10.000  -0.4748   0.12519   0.11770  -0.0154   1.0000   0.2381
  -9.750  -0.4330   0.11848   0.11090  -0.0127   1.0000   0.2496
  -9.500  -0.4556   0.11725   0.10980  -0.0155   1.0000   0.2575
  -9.250  -0.4242   0.11212   0.10462  -0.0138   1.0000   0.2690
  -9.000  -0.4272   0.10913   0.10171  -0.0147   1.0000   0.2790
  -8.750  -0.4384   0.10773   0.10041  -0.0156   1.0000   0.2932
  -8.500  -0.4147   0.10332   0.09599  -0.0142   1.0000   0.3092
  -8.250  -0.4033   0.09993   0.09264  -0.0135   1.0000   0.3251
  -8.000  -0.3958   0.09689   0.08965  -0.0129   1.0000   0.3419
  -7.750  -0.3894   0.09404   0.08686  -0.0122   1.0000   0.3605
  -7.500  -0.3879   0.09170   0.08459  -0.0111   1.0000   0.3823
  -7.250  -0.4105   0.09099   0.08406  -0.0093   1.0000   0.4041
  -7.000  -0.3722   0.08687   0.07989  -0.0070   1.0000   0.4380
  -6.750  -0.3514   0.08393   0.07696  -0.0048   1.0000   0.4716
  -6.500  -0.3350   0.08133   0.07441  -0.0025   1.0000   0.5064
  -6.250  -0.3259   0.07948   0.07262   0.0006   1.0000   0.5455
  -6.000  -0.2948   0.07621   0.06935   0.0022   1.0000   0.5912
  -5.000  -0.5229   0.05437   0.04637  -0.0237   1.0000   0.2193
  -4.750  -0.5037   0.04969   0.04132  -0.0244   1.0000   0.1984
  -4.500  -0.4855   0.04607   0.03743  -0.0245   1.0000   0.1892
  -4.250  -0.4625   0.04340   0.03387  -0.0252   1.0000   0.1816
  -4.000  -0.4413   0.04104   0.03116  -0.0250   1.0000   0.1814
  -3.750  -0.4186   0.03884   0.02862  -0.0250   1.0000   0.1810
  -3.500  -0.3941   0.03713   0.02647  -0.0249   1.0000   0.1822
  -3.250  -0.3713   0.03513   0.02450  -0.0249   1.0000   0.1870
  -3.000  -0.3470   0.03377   0.02293  -0.0247   1.0000   0.1914
  -2.750  -0.3219   0.03263   0.02148  -0.0245   1.0000   0.1970
  -2.500  -0.2974   0.03146   0.02016  -0.0243   1.0000   0.2048
  -2.250  -0.2733   0.03057   0.01919  -0.0239   1.0000   0.2142
  -2.000  -0.2492   0.02970   0.01829  -0.0235   1.0000   0.2264
  -1.750  -0.2253   0.02900   0.01758  -0.0229   1.0000   0.2416
  -1.500  -0.2010   0.02835   0.01700  -0.0224   1.0000   0.2620
  -1.250  -0.1759   0.02771   0.01648  -0.0221   1.0000   0.2920
  -1.000  -0.1492   0.02670   0.01597  -0.0225   1.0000   0.3494
  -0.750  -0.1404   0.02423   0.01632  -0.0145   1.0000   1.0000
  -0.500  -0.1195   0.02461   0.01606  -0.0138   1.0000   1.0000
  -0.250  -0.0996   0.02499   0.01609  -0.0134   1.0000   1.0000
   0.000  -0.0794   0.02542   0.01626  -0.0132   1.0000   1.0000
   0.250  -0.0590   0.02589   0.01651  -0.0131   1.0000   1.0000
   0.500  -0.0384   0.02641   0.01684  -0.0131   1.0000   1.0000
   0.750  -0.0176   0.02697   0.01725  -0.0131   1.0000   1.0000
   1.000   0.0031   0.02758   0.01772  -0.0132   1.0000   1.0000
   1.250   0.0238   0.02823   0.01825  -0.0133   1.0000   1.0000
   1.500   0.0444   0.02892   0.01885  -0.0135   1.0000   1.0000
   1.750   0.0649   0.02966   0.01951  -0.0137   1.0000   1.0000
   2.000   0.0853   0.03045   0.02022  -0.0139   1.0000   1.0000
   2.250   0.1054   0.03127   0.02100  -0.0141   1.0000   1.0000
   2.500   0.1253   0.03215   0.02184  -0.0144   1.0000   1.0000
   2.750   0.1634   0.03355   0.02322  -0.0181   0.9927   1.0000
   3.000   0.2057   0.03510   0.02478  -0.0226   0.9815   1.0000
   3.250   0.2459   0.03661   0.02632  -0.0266   0.9695   1.0000
   3.500   0.2850   0.03811   0.02786  -0.0303   0.9560   1.0000
   3.750   0.3222   0.03949   0.02932  -0.0335   0.9393   1.0000
   4.000   0.3760   0.04095   0.03090  -0.0388   0.9126   1.0000
   4.250   0.4373   0.04163   0.03173  -0.0434   0.8703   1.0000
   4.500   0.5042   0.04137   0.03167  -0.0473   0.8274   1.0000
   4.750   0.5619   0.04060   0.03114  -0.0492   0.7924   1.0000
   5.000   0.6223   0.03887   0.02969  -0.0501   0.7581   1.0000
   5.250   0.6952   0.03506   0.02627  -0.0499   0.7205   1.0000
   5.500   0.7546   0.02969   0.02128  -0.0452   0.6713   1.0000
   5.750   0.7892   0.02390   0.01511  -0.0346   0.4844   1.0000
   6.000   0.7937   0.02580   0.01515  -0.0289   0.3352   1.0000
   6.250   0.8112   0.02781   0.01650  -0.0271   0.2860   1.0000
   6.500   0.8412   0.02959   0.01791  -0.0269   0.2532   1.0000
   6.750   0.8739   0.03129   0.01945  -0.0271   0.2299   1.0000
   7.000   0.9067   0.03312   0.02117  -0.0275   0.2129   1.0000
   7.250   0.9377   0.03516   0.02320  -0.0277   0.2008   1.0000
   7.500   0.9676   0.03735   0.02526  -0.0279   0.1901   1.0000
   7.750   0.9899   0.03948   0.02780  -0.0270   0.1835   1.0000
   8.000   1.0146   0.04173   0.03017  -0.0266   0.1767   1.0000
   8.250   1.0365   0.04451   0.03310  -0.0260   0.1715   1.0000
   8.500   1.0517   0.04726   0.03635  -0.0245   0.1686   1.0000
   8.750   1.0648   0.05021   0.03971  -0.0230   0.1658   1.0000
   9.000   1.0769   0.05321   0.04302  -0.0216   0.1627   1.0000
   9.250   1.0897   0.05640   0.04645  -0.0204   0.1602   1.0000
   9.500   1.0931   0.06022   0.05068  -0.0185   0.1601   1.0000
   9.750   1.0861   0.06453   0.05546  -0.0162   0.1612   1.0000
  10.000   1.0674   0.06940   0.06079  -0.0137   0.1632   1.0000
  10.250   1.0395   0.07472   0.06646  -0.0115   0.1660   1.0000
  10.500   1.0109   0.08003   0.07197  -0.0098   0.1687   1.0000
  10.750   0.9908   0.08576   0.07780  -0.0095   0.1710   1.0000
  11.000   0.9882   0.09148   0.08357  -0.0099   0.1728   1.0000
<< Back to RUTAN CANARD AIRFOIL (amsoil1-il)

Polar data table (+)

Polar graphs


<< Back to RUTAN CANARD AIRFOIL (amsoil1-il)