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

RAF 32 MOD AIRFOIL (raf32md-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: RAF 32 MOD AIRFOIL (raf32md-il)
Reynolds number: 200,000
Max Cl/Cd: 87.25 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf32md-il-200000.txt
Download as CSV file: xf-raf32md-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 32 MOD 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4122   0.09903   0.09571  -0.0414   1.0000   0.0576
  -8.750  -0.4330   0.09715   0.09393  -0.0401   1.0000   0.0578
  -8.500  -0.4534   0.09565   0.09252  -0.0380   1.0000   0.0579
  -8.000  -0.4356   0.09056   0.08743  -0.0303   1.0000   0.0614
  -7.750  -0.4397   0.08857   0.08548  -0.0295   0.9990   0.0629
  -7.500  -0.4288   0.08405   0.08095  -0.0366   0.9932   0.0669
  -7.250  -0.4136   0.07070   0.06755  -0.0622   0.9825   0.0714
  -7.000  -0.3959   0.07010   0.06696  -0.0583   0.9786   0.0732
  -6.750  -0.3663   0.06175   0.05837  -0.0748   0.9718   0.0821
  -6.500  -0.3552   0.05491   0.05137  -0.0804   0.9642   0.0853
  -6.250  -0.3291   0.05370   0.05022  -0.0803   0.9608   0.0886
  -6.000  -0.3111   0.03349   0.02828  -0.0932   0.9531   0.0557
  -5.750  -0.2861   0.02862   0.02280  -0.0946   0.9482   0.0525
  -5.500  -0.2520   0.02586   0.01952  -0.0966   0.9452   0.0535
  -5.250  -0.2148   0.02370   0.01691  -0.0988   0.9432   0.0547
  -5.000  -0.1917   0.02228   0.01517  -0.0980   0.9366   0.0553
  -4.750  -0.1577   0.02118   0.01379  -0.0992   0.9328   0.0572
  -4.500  -0.1210   0.01969   0.01212  -0.1010   0.9305   0.0592
  -4.250  -0.0825   0.01851   0.01090  -0.1032   0.9286   0.0614
  -4.000  -0.0607   0.01796   0.01033  -0.1020   0.9210   0.0637
  -3.750  -0.0246   0.01735   0.00967  -0.1037   0.9177   0.0679
  -3.500   0.0142   0.01653   0.00883  -0.1058   0.9155   0.0722
  -3.250   0.0554   0.01588   0.00819  -0.1084   0.9139   0.0802
  -3.000   0.0759   0.01552   0.00786  -0.1070   0.9054   0.0925
  -2.750   0.1126   0.01442   0.00728  -0.1091   0.9026   0.1996
  -2.500   0.1514   0.01388   0.00704  -0.1114   0.9004   0.3052
  -2.250   0.1773   0.01369   0.00696  -0.1110   0.8937   0.3548
  -2.000   0.2098   0.01334   0.00676  -0.1119   0.8891   0.4071
  -1.750   0.2459   0.01286   0.00652  -0.1134   0.8862   0.4718
  -1.500   0.2709   0.01265   0.00657  -0.1126   0.8795   0.5577
  -1.250   0.3006   0.01236   0.00645  -0.1127   0.8742   0.6272
  -1.000   0.3335   0.01193   0.00623  -0.1131   0.8709   0.6989
  -0.750   0.3542   0.01174   0.00624  -0.1112   0.8628   0.7632
  -0.500   0.3889   0.01128   0.00602  -0.1118   0.8585   0.8571
  -0.250   0.4568   0.01091   0.00567  -0.1200   0.8568   1.0000
   0.000   0.4818   0.01097   0.00563  -0.1194   0.8484   1.0000
   0.250   0.5138   0.01088   0.00545  -0.1200   0.8426   1.0000
   0.500   0.5398   0.01094   0.00544  -0.1196   0.8345   1.0000
   0.750   0.5703   0.01089   0.00531  -0.1199   0.8280   1.0000
   1.000   0.5962   0.01096   0.00533  -0.1194   0.8196   1.0000
   1.250   0.6264   0.01093   0.00525  -0.1196   0.8128   1.0000
   1.500   0.6515   0.01103   0.00532  -0.1189   0.8038   1.0000
   1.750   0.6822   0.01101   0.00524  -0.1193   0.7971   1.0000
   2.000   0.7062   0.01113   0.00536  -0.1184   0.7874   1.0000
   2.250   0.7364   0.01114   0.00533  -0.1186   0.7803   1.0000
   2.500   0.7608   0.01125   0.00544  -0.1178   0.7704   1.0000
   2.750   0.7874   0.01133   0.00554  -0.1174   0.7615   1.0000
   3.000   0.8158   0.01137   0.00556  -0.1173   0.7531   1.0000
   3.250   0.8401   0.01148   0.00569  -0.1164   0.7424   1.0000
   3.500   0.8661   0.01151   0.00572  -0.1158   0.7310   1.0000
   3.750   0.8918   0.01150   0.00571  -0.1149   0.7173   1.0000
   4.000   0.9162   0.01147   0.00567  -0.1138   0.7012   1.0000
   4.250   0.9402   0.01148   0.00566  -0.1127   0.6844   1.0000
   4.500   0.9643   0.01154   0.00573  -0.1116   0.6680   1.0000
   4.750   0.9883   0.01163   0.00579  -0.1105   0.6506   1.0000
   5.000   1.0107   0.01175   0.00591  -0.1091   0.6308   1.0000
   5.250   1.0326   0.01190   0.00606  -0.1077   0.6095   1.0000
   5.500   1.0546   0.01210   0.00625  -0.1063   0.5885   1.0000
   5.750   1.0749   0.01232   0.00648  -0.1046   0.5640   1.0000
   6.000   1.0935   0.01259   0.00671  -0.1026   0.5333   1.0000
   6.250   1.1087   0.01298   0.00696  -0.1000   0.4895   1.0000
   6.500   1.1183   0.01362   0.00731  -0.0964   0.4252   1.0000
   6.750   1.1201   0.01470   0.00791  -0.0917   0.3350   1.0000
   7.000   1.1146   0.01634   0.00886  -0.0862   0.2262   1.0000
   7.250   1.1100   0.01801   0.00997  -0.0810   0.1448   1.0000
   7.500   1.1121   0.01939   0.01103  -0.0769   0.1028   1.0000
   7.750   1.1166   0.02065   0.01214  -0.0733   0.0846   1.0000
   8.000   1.1229   0.02186   0.01331  -0.0700   0.0726   1.0000
   8.250   1.1312   0.02302   0.01451  -0.0671   0.0634   1.0000
   8.500   1.1362   0.02449   0.01601  -0.0639   0.0564   1.0000
   8.750   1.1447   0.02577   0.01729  -0.0614   0.0503   1.0000
   9.000   1.1531   0.02738   0.01892  -0.0589   0.0461   1.0000
   9.250   1.1665   0.02863   0.02025  -0.0570   0.0429   1.0000
   9.500   1.1797   0.03001   0.02162  -0.0553   0.0399   1.0000
   9.750   1.1984   0.03187   0.02351  -0.0542   0.0369   1.0000
  10.000   1.2163   0.03313   0.02492  -0.0529   0.0347   1.0000
  10.250   1.2361   0.03463   0.02652  -0.0520   0.0329   1.0000
  10.500   1.2570   0.03630   0.02826  -0.0513   0.0315   1.0000
  10.750   1.2996   0.04031   0.03239  -0.0538   0.0296   1.0000
  11.000   1.3035   0.04145   0.03380  -0.0507   0.0287   1.0000
  11.250   1.3134   0.04348   0.03613  -0.0487   0.0278   1.0000
  11.500   1.3228   0.04618   0.03915  -0.0468   0.0272   1.0000
  11.750   1.3270   0.04921   0.04250  -0.0445   0.0269   1.0000
  12.000   1.3255   0.05244   0.04605  -0.0418   0.0267   1.0000
  12.250   1.3194   0.05600   0.04994  -0.0391   0.0268   1.0000
  12.500   1.3084   0.05983   0.05408  -0.0365   0.0269   1.0000
  12.750   1.2935   0.06392   0.05847  -0.0342   0.0271   1.0000
  13.000   1.2760   0.06833   0.06316  -0.0324   0.0274   1.0000
  13.250   1.2567   0.07306   0.06815  -0.0312   0.0277   1.0000
  13.500   1.2361   0.07812   0.07343  -0.0309   0.0279   1.0000
  13.750   1.2140   0.08366   0.07919  -0.0313   0.0282   1.0000
  14.000   1.1918   0.08960   0.08532  -0.0326   0.0284   1.0000
  14.250   1.1698   0.09605   0.09194  -0.0347   0.0288   1.0000
  14.500   1.1466   0.10315   0.09920  -0.0376   0.0290   1.0000
<< Back to RAF 32 MOD AIRFOIL (raf32md-il)

Polar data table (+)

Polar graphs


<< Back to RAF 32 MOD AIRFOIL (raf32md-il)