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

20-32C AIRFOIL (2032c-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: 20-32C AIRFOIL (2032c-il)
Reynolds number: 100,000
Max Cl/Cd: 54.9 at α=9°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-2032c-il-100000.txt
Download as CSV file: xf-2032c-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -6.750  -0.2228   0.09488   0.09094  -0.0196   1.0000   0.0464
  -6.500  -0.2283   0.09353   0.08970  -0.0179   1.0000   0.0473
  -6.250  -0.2377   0.09261   0.08889  -0.0157   1.0000   0.0481
  -6.000  -0.2490   0.09191   0.08831  -0.0135   1.0000   0.0489
  -5.750  -0.2600   0.09123   0.08773  -0.0116   1.0000   0.0497
  -5.500  -0.2083   0.08847   0.08491  -0.0292   0.9901   0.0544
  -5.250  -0.1599   0.08256   0.07895  -0.0422   0.9810   0.0560
  -5.000  -0.1375   0.07703   0.07342  -0.0421   0.9731   0.0585
  -4.750  -0.0942   0.07244   0.06878  -0.0504   0.9630   0.0633
  -4.500  -0.0098   0.06760   0.06371  -0.0736   0.9525   0.0695
  -4.250   0.0133   0.06241   0.05857  -0.0732   0.9439   0.0730
  -4.000   0.0920   0.05827   0.05415  -0.0915   0.9320   0.0838
  -3.750   0.1108   0.05388   0.04984  -0.0899   0.9207   0.0898
  -3.500   0.1645   0.04981   0.04558  -0.0994   0.9074   0.1005
  -3.250   0.2145   0.04655   0.04209  -0.1070   0.8922   0.1135
  -3.000   0.2384   0.04354   0.03907  -0.1068   0.8756   0.1219
  -2.750   0.2849   0.04100   0.03619  -0.1129   0.8583   0.1423
  -2.500   0.3105   0.03830   0.03341  -0.1132   0.8410   0.1580
  -2.250   0.3379   0.03600   0.03097  -0.1137   0.8219   0.1741
  -1.750   0.3942   0.03216   0.02679  -0.1146   0.7833   0.2232
  -1.500   0.4207   0.03040   0.02486  -0.1146   0.7631   0.2611
  -0.750   0.5597   0.02209   0.01399  -0.1227   0.7068   0.1110
  -0.500   0.5892   0.02076   0.01232  -0.1224   0.6861   0.1075
  -0.250   0.6184   0.01981   0.01092  -0.1219   0.6670   0.1062
   0.000   0.6466   0.01919   0.00997  -0.1213   0.6464   0.1089
   0.250   0.6739   0.01868   0.00922  -0.1206   0.6265   0.1154
   0.500   0.7011   0.01829   0.00862  -0.1200   0.6083   0.1206
   0.750   0.7280   0.01796   0.00816  -0.1193   0.5889   0.1290
   1.000   0.7550   0.01772   0.00783  -0.1187   0.5708   0.1490
   1.250   0.7778   0.01581   0.00753  -0.1172   0.5551   1.0000
   1.500   0.8048   0.01616   0.00742  -0.1167   0.5388   1.0000
   1.750   0.8316   0.01654   0.00748  -0.1162   0.5235   1.0000
   2.000   0.8584   0.01694   0.00762  -0.1158   0.5092   1.0000
   2.250   0.8851   0.01736   0.00784  -0.1155   0.4959   1.0000
   2.500   0.9117   0.01780   0.00810  -0.1152   0.4835   1.0000
   2.750   0.9384   0.01825   0.00840  -0.1149   0.4722   1.0000
   3.000   0.9652   0.01872   0.00864  -0.1147   0.4625   1.0000
   3.250   0.9916   0.01920   0.00907  -0.1144   0.4522   1.0000
   3.500   1.0181   0.01973   0.00953  -0.1142   0.4434   1.0000
   3.750   1.0449   0.02023   0.00991  -0.1140   0.4356   1.0000
   4.000   1.0708   0.02079   0.01050  -0.1138   0.4270   1.0000
   4.250   1.0977   0.02133   0.01087  -0.1136   0.4204   1.0000
   4.500   1.1232   0.02195   0.01161  -0.1134   0.4128   1.0000
   4.750   1.1499   0.02253   0.01214  -0.1132   0.4070   1.0000
   5.000   1.1752   0.02324   0.01294  -0.1130   0.4006   1.0000
   5.250   1.2011   0.02384   0.01356  -0.1127   0.3945   1.0000
   5.500   1.2271   0.02455   0.01426  -0.1126   0.3892   1.0000
   5.750   1.2515   0.02531   0.01520  -0.1123   0.3833   1.0000
   6.000   1.2776   0.02601   0.01594  -0.1121   0.3788   1.0000
   6.250   1.3023   0.02691   0.01697  -0.1118   0.3741   1.0000
   6.500   1.3259   0.02772   0.01798  -0.1114   0.3683   1.0000
   6.750   1.3506   0.02794   0.01816  -0.1108   0.3592   1.0000
   7.000   1.3756   0.02789   0.01806  -0.1102   0.3488   1.0000
   7.250   1.3974   0.02777   0.01799  -0.1092   0.3358   1.0000
   7.500   1.4188   0.02779   0.01811  -0.1082   0.3239   1.0000
   7.750   1.4410   0.02796   0.01837  -0.1073   0.3140   1.0000
   8.000   1.4619   0.02782   0.01831  -0.1062   0.3013   1.0000
   8.250   1.4800   0.02759   0.01820  -0.1046   0.2857   1.0000
   8.500   1.4987   0.02764   0.01840  -0.1032   0.2727   1.0000
   8.750   1.5133   0.02757   0.01846  -0.1013   0.2514   1.0000
   9.000   1.5289   0.02785   0.01890  -0.0996   0.2273   1.0000
   9.250   1.5434   0.02853   0.01960  -0.0979   0.1934   1.0000
   9.500   1.5403   0.03103   0.02143  -0.0949   0.1009   1.0000
   9.750   1.5250   0.03474   0.02475  -0.0908   0.0510   1.0000
  10.000   1.5209   0.03703   0.02715  -0.0875   0.0459   1.0000
  10.250   1.5160   0.03957   0.02987  -0.0848   0.0430   1.0000
  10.500   1.5086   0.04258   0.03307  -0.0827   0.0411   1.0000
  10.750   1.5022   0.04577   0.03653  -0.0814   0.0398   1.0000
  11.000   1.4927   0.04964   0.04066  -0.0809   0.0388   1.0000
  11.250   1.4801   0.05430   0.04557  -0.0812   0.0381   1.0000
  11.500   1.4647   0.05979   0.05130  -0.0823   0.0376   1.0000
  11.750   1.4465   0.06611   0.05786  -0.0842   0.0374   1.0000
  12.000   1.4263   0.07306   0.06504  -0.0865   0.0375   1.0000
  12.250   1.4053   0.08038   0.07257  -0.0889   0.0377   1.0000
  12.500   1.3847   0.08771   0.08009  -0.0914   0.0379   1.0000
  12.750   1.3650   0.09494   0.08750  -0.0937   0.0379   1.0000
<< Back to 20-32C AIRFOIL (2032c-il)

Polar data table (+)

Polar graphs


<< Back to 20-32C AIRFOIL (2032c-il)