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RAF 26 AIRFOIL (raf26-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: RAF 26 AIRFOIL (raf26-il)
Reynolds number: 50,000
Max Cl/Cd: 33.96 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf26-il-50000.txt
Download as CSV file: xf-raf26-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 26 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4879   0.09609   0.08928  -0.0087   1.0000   0.2499
  -7.500  -0.5066   0.09528   0.08863  -0.0090   1.0000   0.2597
  -7.250  -0.5040   0.09204   0.08547  -0.0080   1.0000   0.2735
  -6.750  -0.4881   0.08505   0.07858  -0.0040   1.0000   0.3134
  -6.500  -0.4785   0.08156   0.07508  -0.0016   1.0000   0.3358
  -6.250  -0.4917   0.08007   0.07372  -0.0015   1.0000   0.3571
  -6.000  -0.4748   0.07600   0.06968   0.0021   1.0000   0.3823
  -5.750  -0.4678   0.07270   0.06643   0.0045   1.0000   0.4049
  -5.500  -0.4649   0.06986   0.06367   0.0070   1.0000   0.4310
  -4.750  -0.4029   0.04199   0.03381  -0.0414   1.0000   0.1334
  -4.500  -0.3796   0.03760   0.02846  -0.0417   1.0000   0.1213
  -4.250  -0.3581   0.03411   0.02466  -0.0409   1.0000   0.1189
  -4.000  -0.3349   0.03122   0.02119  -0.0400   1.0000   0.1196
  -3.750  -0.3128   0.02885   0.01859  -0.0390   1.0000   0.1282
  -3.500  -0.2876   0.02662   0.01582  -0.0378   1.0000   0.1336
  -3.250  -0.2629   0.02453   0.01354  -0.0367   1.0000   0.1436
  -3.000  -0.2386   0.02286   0.01181  -0.0356   1.0000   0.1689
  -2.750  -0.2122   0.02111   0.00997  -0.0345   1.0000   0.2053
  -2.500  -0.1861   0.01874   0.00853  -0.0337   1.0000   0.3384
  -2.250  -0.1343   0.01554   0.00733  -0.0347   1.0000   1.0000
  -2.000  -0.1126   0.01554   0.00679  -0.0339   1.0000   1.0000
  -1.750  -0.0913   0.01558   0.00640  -0.0330   1.0000   1.0000
  -1.500  -0.0701   0.01565   0.00607  -0.0321   1.0000   1.0000
  -1.250  -0.0491   0.01575   0.00586  -0.0312   1.0000   1.0000
  -1.000  -0.0282   0.01588   0.00572  -0.0302   1.0000   1.0000
  -0.750  -0.0075   0.01603   0.00565  -0.0293   1.0000   1.0000
  -0.500   0.0131   0.01621   0.00563  -0.0284   1.0000   1.0000
  -0.250   0.0335   0.01642   0.00564  -0.0275   1.0000   1.0000
   0.000   0.0537   0.01665   0.00573  -0.0266   1.0000   1.0000
   0.250   0.0738   0.01692   0.00588  -0.0257   1.0000   1.0000
   0.500   0.0936   0.01721   0.00608  -0.0249   1.0000   1.0000
   0.750   0.1133   0.01753   0.00633  -0.0241   1.0000   1.0000
   1.000   0.1328   0.01789   0.00663  -0.0233   1.0000   1.0000
   1.250   0.1520   0.01829   0.00699  -0.0226   1.0000   1.0000
   1.500   0.1710   0.01872   0.00740  -0.0218   1.0000   1.0000
   1.750   0.1897   0.01918   0.00787  -0.0212   1.0000   1.0000
   2.000   0.2082   0.01969   0.00840  -0.0205   1.0000   1.0000
   2.250   0.2265   0.02025   0.00902  -0.0200   1.0000   1.0000
   2.500   0.2444   0.02085   0.00968  -0.0194   1.0000   1.0000
   2.750   0.2619   0.02151   0.01041  -0.0190   1.0000   1.0000
   3.000   0.2791   0.02224   0.01122  -0.0186   1.0000   1.0000
   3.250   0.2957   0.02303   0.01212  -0.0183   1.0000   1.0000
   3.500   0.3117   0.02392   0.01314  -0.0181   1.0000   1.0000
   3.750   0.3476   0.02522   0.01472  -0.0220   0.9896   1.0000
   4.000   0.4534   0.02651   0.01664  -0.0373   0.9395   1.0000
   4.250   0.6042   0.02409   0.01550  -0.0535   0.8613   1.0000
   4.500   0.6876   0.02051   0.01290  -0.0541   0.7721   1.0000
   4.750   0.6986   0.02057   0.01085  -0.0415   0.3457   1.0000
   5.000   0.7100   0.02376   0.01270  -0.0386   0.2209   1.0000
   5.250   0.7368   0.02683   0.01520  -0.0379   0.1587   1.0000
   5.500   0.7701   0.02965   0.01799  -0.0376   0.1297   1.0000
   5.750   0.8000   0.03238   0.02091  -0.0370   0.1149   1.0000
   6.000   0.8291   0.03550   0.02434  -0.0361   0.1092   1.0000
   6.250   0.8531   0.03897   0.02813  -0.0349   0.1048   1.0000
   6.500   0.8718   0.04202   0.03192  -0.0328   0.1011   1.0000
   6.750   0.8887   0.04575   0.03620  -0.0308   0.1015   1.0000
   7.000   0.9042   0.05008   0.04096  -0.0289   0.1044   1.0000
   7.250   0.9122   0.05447   0.04617  -0.0263   0.1117   1.0000
   7.500   0.9267   0.06006   0.05193  -0.0252   0.1187   1.0000
   7.750   0.9205   0.06533   0.05812  -0.0231   0.1358   1.0000
   8.000   0.9245   0.07418   0.06742  -0.0242   0.1737   1.0000
   8.250   0.8787   0.08144   0.07507  -0.0282   0.2051   1.0000
   8.500   0.7956   0.07554   0.06949  -0.0188   0.1755   1.0000
   8.750   0.7558   0.08170   0.07561  -0.0213   0.1776   1.0000
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