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20-32C AIRFOIL (2032c-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: 20-32C AIRFOIL (2032c-il)
Reynolds number: 50,000
Max Cl/Cd: 31.7 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-2032c-il-50000.txt
Download as CSV file: xf-2032c-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: 20-32C 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.2446   0.11407   0.10788  -0.0201   1.0000   0.0956
  -8.000  -0.2475   0.11417   0.10812  -0.0218   1.0000   0.0969
  -7.750  -0.2530   0.11491   0.10902  -0.0240   1.0000   0.0975
  -7.500  -0.2326   0.10640   0.10052  -0.0214   1.0000   0.1007
  -7.250  -0.2270   0.10345   0.09765  -0.0208   1.0000   0.1042
  -7.000  -0.2270   0.10172   0.09604  -0.0205   1.0000   0.1074
  -6.750  -0.2298   0.10096   0.09544  -0.0212   1.0000   0.1103
  -6.500  -0.2352   0.10161   0.09625  -0.0237   1.0000   0.1117
  -6.250  -0.2376   0.10049   0.09527  -0.0248   1.0000   0.1126
  -6.000  -0.2347   0.09499   0.08985  -0.0189   1.0000   0.1159
  -5.750  -0.2378   0.09320   0.08817  -0.0171   1.0000   0.1190
  -5.500  -0.2425   0.09209   0.08716  -0.0165   1.0000   0.1224
  -5.250  -0.2448   0.09223   0.08740  -0.0194   1.0000   0.1257
  -4.750  -0.2495   0.08760   0.08296  -0.0165   1.0000   0.1307
  -4.500  -0.2491   0.08593   0.08136  -0.0163   1.0000   0.1361
  -4.250  -0.2347   0.08537   0.08078  -0.0238   1.0000   0.1420
  -4.000  -0.2399   0.08224   0.07777  -0.0193   1.0000   0.1460
  -3.750  -0.2119   0.08076   0.07624  -0.0284   0.9978   0.1567
  -3.500  -0.1606   0.07544   0.07086  -0.0374   0.9845   0.1722
  -3.250  -0.1097   0.07057   0.06593  -0.0457   0.9703   0.1906
  -3.000  -0.0468   0.06613   0.06135  -0.0577   0.9549   0.2171
  -2.750   0.0020   0.06206   0.05719  -0.0647   0.9387   0.2506
  -2.500   0.0517   0.05810   0.05317  -0.0716   0.9220   0.2936
  -2.250   0.0979   0.05414   0.04917  -0.0769   0.9059   0.3480
  -1.500   0.1619   0.04319   0.03847  -0.0674   0.8612   0.6068
  -1.000   0.4963   0.03560   0.02784  -0.1339   0.8226   0.1971
  -0.750   0.5400   0.03345   0.02515  -0.1356   0.8029   0.1801
  -0.500   0.5796   0.03183   0.02304  -0.1363   0.7844   0.1823
  -0.250   0.6150   0.03048   0.02121  -0.1360   0.7658   0.1822
   0.000   0.6454   0.02963   0.01993  -0.1352   0.7447   0.1842
   0.250   0.6759   0.02875   0.01871  -0.1339   0.7262   0.1919
   0.500   0.7044   0.02821   0.01787  -0.1325   0.7078   0.2114
   0.750   0.7304   0.02789   0.01743  -0.1312   0.6880   0.2430
   1.000   0.7575   0.02681   0.01666  -0.1299   0.6710   0.3477
   1.250   0.7812   0.02549   0.01587  -0.1275   0.6554   1.0000
   1.500   0.8074   0.02596   0.01592  -0.1264   0.6392   1.0000
   1.750   0.8333   0.02651   0.01613  -0.1255   0.6238   1.0000
   2.000   0.8588   0.02713   0.01649  -0.1246   0.6090   1.0000
   2.250   0.8838   0.02788   0.01702  -0.1240   0.5950   1.0000
   2.500   0.9084   0.02871   0.01769  -0.1234   0.5815   1.0000
   2.750   0.9327   0.02963   0.01849  -0.1228   0.5688   1.0000
   3.000   0.9566   0.03064   0.01942  -0.1224   0.5572   1.0000
   3.250   0.9819   0.03148   0.02013  -0.1218   0.5475   1.0000
   3.500   1.0062   0.03245   0.02103  -0.1213   0.5374   1.0000
   3.750   1.0274   0.03392   0.02253  -0.1211   0.5270   1.0000
   4.000   1.0523   0.03485   0.02342  -0.1205   0.5190   1.0000
   4.250   1.0728   0.03648   0.02512  -0.1203   0.5102   1.0000
   4.500   1.0934   0.03811   0.02681  -0.1200   0.5025   1.0000
   4.750   1.1149   0.03954   0.02826  -0.1196   0.4948   1.0000
   5.000   1.1311   0.04178   0.03063  -0.1193   0.4873   1.0000
   5.250   1.1500   0.04365   0.03264  -0.1190   0.4808   1.0000
   5.500   1.1650   0.04615   0.03526  -0.1188   0.4750   1.0000
   5.750   1.1687   0.05007   0.03941  -0.1188   0.4682   1.0000
   6.000   1.2002   0.05025   0.03955  -0.1179   0.4635   1.0000
   6.250   1.1521   0.06111   0.05079  -0.1194   0.4570   1.0000
   6.500   1.0670   0.07694   0.06674  -0.1241   0.4575   1.0000
   6.750   1.0401   0.08482   0.07461  -0.1262   0.4593   1.0000
   7.000   1.0310   0.09057   0.08040  -0.1274   0.4611   1.0000
   7.250   1.0388   0.09486   0.08479  -0.1282   0.4631   1.0000
   7.500   0.9081   0.11777   0.10789  -0.1434   0.5936   1.0000
   7.750   0.8981   0.11965   0.10978  -0.1415   0.5823   1.0000
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