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

RAF 19 AIRFOIL (raf19-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: RAF 19 AIRFOIL (raf19-il)
Reynolds number: 200,000
Max Cl/Cd: 61.76 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-raf19-il-200000-n5.txt
Download as CSV file: xf-raf19-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 19 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500   0.0382   0.10939   0.10614  -0.0700   0.9358   0.0257
  -9.250   0.0495   0.10655   0.10331  -0.0715   0.9239   0.0258
  -9.000   0.0578   0.10140   0.09814  -0.0753   0.9130   0.0192
  -8.750   0.0796   0.09821   0.09493  -0.0785   0.9042   0.0185
  -8.500   0.1000   0.09454   0.09123  -0.0825   0.8913   0.0180
  -8.250   0.1264   0.09040   0.08704  -0.0884   0.8784   0.0177
  -8.000   0.1614   0.08572   0.08227  -0.0967   0.8635   0.0179
  -7.500   0.2268   0.07635   0.07260  -0.1130   0.8161   0.0187
  -7.250   0.2426   0.07321   0.06929  -0.1162   0.7920   0.0187
  -7.000   0.2510   0.07088   0.06685  -0.1172   0.7706   0.0187
  -6.750   0.2580   0.06822   0.06407  -0.1184   0.7517   0.0189
  -6.500   0.2650   0.06548   0.06125  -0.1197   0.7361   0.0191
  -6.250   0.2773   0.06391   0.05961  -0.1204   0.7213   0.0199
  -6.000   0.2919   0.06287   0.05849  -0.1207   0.7079   0.0215
  -5.750   0.3035   0.06048   0.05603  -0.1222   0.6961   0.0220
  -5.500   0.3155   0.05745   0.05297  -0.1245   0.6847   0.0233
  -5.250   0.3301   0.05431   0.04977  -0.1272   0.6733   0.0245
  -5.000   0.3482   0.05308   0.04846  -0.1280   0.6603   0.0258
  -4.750   0.3680   0.05002   0.04532  -0.1314   0.6481   0.0298
  -4.500   0.3880   0.04888   0.04410  -0.1322   0.6343   0.0324
  -4.250   0.4181   0.04466   0.03970  -0.1389   0.6224   0.0418
  -4.000   0.4357   0.04477   0.03973  -0.1373   0.6080   0.0440
  -3.750   0.4655   0.04277   0.03751  -0.1411   0.5952   0.0507
  -3.250   0.5152   0.04167   0.03613  -0.1431   0.5708   0.0635
  -1.750   0.6686   0.03121   0.02458  -0.1502   0.5159   0.0640
  -1.500   0.6937   0.02963   0.02286  -0.1507   0.5075   0.0644
  -1.250   0.7159   0.02876   0.02185  -0.1503   0.4990   0.0652
  -1.000   0.7393   0.02792   0.02091  -0.1500   0.4902   0.0657
  -0.750   0.7644   0.02640   0.01916  -0.1501   0.4823   0.0656
  -0.500   0.7887   0.02524   0.01786  -0.1498   0.4739   0.0660
  -0.250   0.8121   0.02422   0.01664  -0.1492   0.4659   0.0665
   0.000   0.8356   0.02324   0.01550  -0.1486   0.4565   0.0673
   0.250   0.8585   0.02230   0.01435  -0.1478   0.4473   0.0685
   0.500   0.8815   0.02096   0.01274  -0.1469   0.4376   0.0703
   0.750   0.9068   0.01791   0.00894  -0.1466   0.4291   0.0773
   1.000   0.9278   0.01803   0.00900  -0.1451   0.4189   0.0788
   1.250   0.9501   0.01798   0.00886  -0.1440   0.4087   0.0809
   1.500   0.9723   0.01770   0.00836  -0.1428   0.3993   0.0853
   1.750   0.9943   0.01758   0.00807  -0.1416   0.3900   0.0892
   2.000   1.0147   0.01782   0.00829  -0.1401   0.3812   0.0920
   2.250   1.0346   0.01797   0.00832  -0.1385   0.3729   0.0980
   2.500   1.0550   0.01809   0.00835  -0.1370   0.3654   0.1035
   2.750   1.0733   0.01838   0.00861  -0.1351   0.3583   0.1071
   3.000   1.0929   0.01854   0.00869  -0.1335   0.3520   0.1122
   3.250   1.1137   0.01862   0.00864  -0.1321   0.3462   0.1172
   3.500   1.1320   0.01891   0.00891  -0.1304   0.3411   0.1201
   3.750   1.1516   0.01918   0.00918  -0.1289   0.3366   0.1238
   4.000   1.1720   0.01944   0.00938  -0.1275   0.3321   0.1304
   4.250   1.1913   0.01972   0.00966  -0.1260   0.3278   0.1342
   4.500   1.2099   0.02006   0.00996  -0.1244   0.3241   0.1372
   4.750   1.2312   0.02026   0.01015  -0.1233   0.3203   0.1398
   5.000   1.2520   0.02049   0.01037  -0.1221   0.3162   0.1433
   5.250   1.2718   0.02077   0.01059  -0.1208   0.3122   0.1463
   5.500   1.2914   0.02106   0.01086  -0.1195   0.3089   0.1482
   5.750   1.3119   0.02138   0.01120  -0.1185   0.3059   0.1504
   6.000   1.3355   0.02167   0.01154  -0.1180   0.3028   0.1528
   6.250   1.3591   0.02201   0.01191  -0.1176   0.2995   0.1558
   6.500   1.3822   0.02238   0.01229  -0.1172   0.2965   0.1594
   6.750   1.4046   0.02280   0.01270  -0.1167   0.2938   0.1626
   7.000   1.4266   0.02327   0.01317  -0.1161   0.2913   0.1660
   7.250   1.4482   0.02366   0.01365  -0.1154   0.2889   0.1713
   7.500   1.4688   0.02406   0.01411  -0.1146   0.2861   0.1769
   7.750   1.4892   0.02448   0.01462  -0.1138   0.2834   0.1817
   8.000   1.5065   0.02495   0.01512  -0.1125   0.2799   0.1878
   8.250   1.5215   0.02554   0.01569  -0.1109   0.2757   0.1933
   8.500   1.5360   0.02602   0.01629  -0.1092   0.2700   0.1994
   8.750   1.5491   0.02661   0.01693  -0.1075   0.2643   0.2062
   9.000   1.5605   0.02735   0.01765  -0.1056   0.2597   0.2135
   9.250   1.5745   0.02797   0.01839  -0.1041   0.2539   0.2232
   9.500   1.5843   0.02880   0.01925  -0.1022   0.2468   0.2332
   9.750   1.5938   0.02971   0.02023  -0.1004   0.2391   0.2436
  10.000   1.6010   0.03082   0.02136  -0.0985   0.2299   0.2551
  10.250   1.6093   0.03197   0.02256  -0.0969   0.2193   0.2739
  10.750   1.6447   0.03768   0.02857  -0.1035   0.1537   1.0000
  11.000   1.6192   0.04226   0.03291  -0.1005   0.1178   1.0000
  11.250   1.5893   0.04770   0.03820  -0.0981   0.0842   1.0000
  11.500   1.5664   0.05280   0.04327  -0.0966   0.0593   1.0000
  11.750   1.5423   0.05842   0.04891  -0.0956   0.0318   1.0000
  12.000   1.5267   0.06330   0.05388  -0.0951   0.0208   1.0000
  12.250   1.5192   0.06733   0.05803  -0.0949   0.0187   1.0000
  12.500   1.5133   0.07127   0.06212  -0.0948   0.0176   1.0000
  12.750   1.5069   0.07535   0.06634  -0.0948   0.0168   1.0000
  13.000   1.4997   0.07963   0.07078  -0.0950   0.0163   1.0000
  13.250   1.4913   0.08416   0.07547  -0.0953   0.0158   1.0000
  13.500   1.4818   0.08895   0.08042  -0.0958   0.0154   1.0000
  13.750   1.4702   0.09413   0.08576  -0.0964   0.0149   1.0000
  14.000   1.4574   0.09959   0.09139  -0.0972   0.0145   1.0000
  14.250   1.4443   0.10521   0.09717  -0.0982   0.0142   1.0000
  14.500   1.4311   0.11093   0.10306  -0.0994   0.0140   1.0000
  14.750   1.4170   0.11688   0.10916  -0.1007   0.0137   1.0000
<< Back to RAF 19 AIRFOIL (raf19-il)

Polar data table (+)

Polar graphs


<< Back to RAF 19 AIRFOIL (raf19-il)