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

RAF 69 AIRFOIL (raf69-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: RAF 69 AIRFOIL (raf69-il)
Reynolds number: 50,000
Max Cl/Cd: 4.71 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf69-il-50000.txt
Download as CSV file: xf-raf69-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 69 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
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.2417   0.12192   0.11360  -0.0279   1.0000   0.4262
 -10.500  -0.2324   0.11989   0.11161  -0.0266   1.0000   0.4373
 -10.250  -0.2547   0.11934   0.11116  -0.0241   1.0000   0.4442
 -10.000  -0.2342   0.11657   0.10840  -0.0228   1.0000   0.4546
  -9.750  -0.2792   0.11831   0.11028  -0.0181   1.0000   0.4622
  -9.500  -0.2479   0.11421   0.10617  -0.0175   1.0000   0.4703
  -9.250  -0.2768   0.11499   0.10705  -0.0132   1.0000   0.4811
  -9.000  -0.2738   0.11261   0.10471  -0.0113   1.0000   0.4863
  -8.750  -0.2722   0.11148   0.10361  -0.0090   1.0000   0.4965
  -8.500  -0.3173   0.11237   0.10461  -0.0041   1.0000   0.5022
  -8.250  -0.6701   0.07448   0.06689  -0.0133   1.0000   0.3442
  -8.000  -0.7760   0.06656   0.05899  -0.0055   1.0000   0.3419
  -7.750  -0.8635   0.05908   0.05129   0.0021   1.0000   0.3424
  -7.500  -0.8873   0.05583   0.04789   0.0064   1.0000   0.3478
  -7.250  -0.8629   0.05654   0.04871   0.0080   1.0000   0.3569
  -7.000  -0.9038   0.05176   0.04354   0.0133   1.0000   0.3630
  -6.750  -0.8808   0.05216   0.04409   0.0147   1.0000   0.3715
  -6.500  -0.8896   0.04992   0.04165   0.0180   1.0000   0.3800
  -6.250  -0.8857   0.04859   0.04024   0.0205   1.0000   0.3883
  -6.000  -0.8774   0.04781   0.03941   0.0227   1.0000   0.3976
  -5.750  -0.8766   0.04615   0.03759   0.0253   1.0000   0.4066
  -5.500  -0.8634   0.04579   0.03725   0.0271   1.0000   0.4161
  -5.250  -0.8600   0.04435   0.03564   0.0294   1.0000   0.4261
  -5.000  -0.8466   0.04401   0.03532   0.0312   1.0000   0.4360
  -4.750  -0.8396   0.04293   0.03411   0.0332   1.0000   0.4463
  -4.500  -0.8274   0.04250   0.03367   0.0349   1.0000   0.4572
  -4.250  -0.8167   0.04184   0.03297   0.0367   1.0000   0.4678
  -4.000  -0.7793   0.04207   0.03307   0.0335   0.9882   0.4845
  -3.750  -0.7425   0.04254   0.03348   0.0308   0.9764   0.5010
  -3.500  -0.7089   0.04307   0.03410   0.0293   0.9661   0.5156
  -3.250  -0.6829   0.04319   0.03422   0.0288   0.9559   0.5310
  -3.000  -0.6507   0.04363   0.03462   0.0272   0.9450   0.5489
  -2.750  -0.6338   0.04349   0.03445   0.0281   0.9368   0.5657
  -2.500  -0.6047   0.04398   0.03494   0.0273   0.9262   0.5850
  -2.250  -0.5852   0.04421   0.03517   0.0281   0.9173   0.6034
  -2.000  -0.5572   0.04490   0.03590   0.0279   0.9071   0.6230
  -1.750  -0.5375   0.04535   0.03642   0.0290   0.8985   0.6412
  -1.500  -0.5107   0.04608   0.03716   0.0291   0.8880   0.6630
  -1.250  -0.4960   0.04648   0.03761   0.0310   0.8804   0.6821
  -1.000  -0.4655   0.04740   0.03852   0.0305   0.8700   0.7065
  -0.750  -0.4595   0.04756   0.03870   0.0336   0.8634   0.7265
  -0.500  -0.4143   0.04905   0.04021   0.0315   0.8511   0.7510
  -0.250  -0.4133   0.04908   0.04027   0.0351   0.8457   0.7716
   0.000  -0.3874   0.04987   0.04109   0.0353   0.8368   0.7950
   0.250  -0.3534   0.05098   0.04219   0.0342   0.8261   0.8195
   0.500  -0.3194   0.05210   0.04331   0.0325   0.8172   0.8434
   0.750  -0.2660   0.05380   0.04498   0.0278   0.8037   0.8664
   1.000  -0.1776   0.05631   0.04743   0.0175   0.7870   0.8866
   1.250  -0.0570   0.05911   0.05014   0.0017   0.7694   0.9028
   1.500  -0.0213   0.06048   0.05152  -0.0026   0.7583   0.9195
   1.750   0.0570   0.06207   0.05305  -0.0126   0.7445   0.9374
   2.000   0.1510   0.06320   0.05414  -0.0246   0.7307   0.9554
   2.250   0.1821   0.06460   0.05557  -0.0293   0.7187   0.9723
   2.500   0.2646   0.06548   0.05642  -0.0404   0.7053   0.9893
   2.750   0.3115   0.06612   0.05706  -0.0465   0.6935   1.0000
   3.000   0.3044   0.06665   0.05754  -0.0438   0.6863   1.0000
   3.250   0.2719   0.06752   0.05837  -0.0385   0.6799   1.0000
   3.500   0.2891   0.06747   0.05825  -0.0379   0.6715   1.0000
   3.750   0.2583   0.06829   0.05903  -0.0325   0.6658   1.0000
   4.000   0.2265   0.06899   0.05967  -0.0271   0.6616   1.0000
   4.250   0.2146   0.06937   0.05999  -0.0233   0.6557   1.0000
   4.500   0.2187   0.06980   0.06033  -0.0207   0.6474   1.0000
   4.750   0.1915   0.07081   0.06125  -0.0158   0.6445   1.0000
   5.000   0.1800   0.07207   0.06242  -0.0128   0.6408   1.0000
   5.250   0.1503   0.07580   0.06612  -0.0111   0.6725   1.0000
   5.500   0.1683   0.07823   0.06849  -0.0115   0.6701   1.0000
   5.750   0.1863   0.07741   0.06757  -0.0090   0.6339   1.0000
   6.000   0.1984   0.07966   0.06978  -0.0089   0.6319   1.0000
   6.250   0.2212   0.08030   0.07035  -0.0079   0.6094   1.0000
   6.500   0.2313   0.08309   0.07313  -0.0081   0.6112   1.0000
   6.750   0.2537   0.08386   0.07385  -0.0073   0.5915   1.0000
   7.000   0.1612   0.08925   0.07930  -0.0043   0.6636   1.0000
   7.250   0.1972   0.09250   0.08253  -0.0061   0.6508   1.0000
   7.500   0.1833   0.09307   0.08307  -0.0036   0.6433   1.0000
   7.750   0.2177   0.09635   0.08634  -0.0053   0.6315   1.0000
   8.000   0.2053   0.09706   0.08703  -0.0031   0.6229   1.0000
   8.250   0.2402   0.10055   0.09052  -0.0049   0.6120   1.0000
   8.500   0.2272   0.10133   0.09129  -0.0028   0.6037   1.0000
   8.750   0.2639   0.10507   0.09504  -0.0046   0.5924   1.0000
   9.000   0.2484   0.10566   0.09561  -0.0026   0.5838   1.0000
   9.250   0.2863   0.10969   0.09966  -0.0044   0.5731   1.0000
   9.500   0.2693   0.11013   0.10009  -0.0025   0.5638   1.0000
   9.750   0.3072   0.11435   0.10433  -0.0043   0.5540   1.0000
  10.000   0.2896   0.11480   0.10476  -0.0025   0.5448   1.0000
  10.250   0.3287   0.11928   0.10928  -0.0043   0.5348   1.0000
  10.500   0.3093   0.11959   0.10958  -0.0027   0.5258   1.0000
  10.750   0.3438   0.12376   0.11377  -0.0041   0.5164   1.0000
<< Back to RAF 69 AIRFOIL (raf69-il)

Polar data table (+)

Polar graphs


<< Back to RAF 69 AIRFOIL (raf69-il)