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=500,000 Ncrit=9


Details Polar file
Airfoil: 20-32C AIRFOIL (2032c-il)
Reynolds number: 500,000
Max Cl/Cd: 103.72 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-2032c-il-500000.txt
Download as CSV file: xf-2032c-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.2463   0.10598   0.10395  -0.0187   1.0000   0.0126
  -8.250  -0.2404   0.10344   0.10144  -0.0209   1.0000   0.0127
  -8.000  -0.2356   0.10095   0.09898  -0.0227   1.0000   0.0127
  -7.750  -0.2269   0.09632   0.09438  -0.0220   1.0000   0.0129
  -7.500  -0.2203   0.09357   0.09166  -0.0217   1.0000   0.0131
  -7.250  -0.2122   0.09108   0.08920  -0.0226   0.9970   0.0135
  -7.000  -0.1886   0.08744   0.08556  -0.0278   0.9834   0.0142
  -6.750  -0.1658   0.08378   0.08188  -0.0330   0.9651   0.0149
  -6.500  -0.1454   0.08028   0.07835  -0.0376   0.9431   0.0156
  -6.250  -0.1270   0.07711   0.07513  -0.0416   0.9183   0.0163
  -6.000  -0.1017   0.07398   0.07191  -0.0489   0.8936   0.0168
  -5.750  -0.0751   0.07050   0.06832  -0.0558   0.8709   0.0169
  -5.500  -0.0495   0.06691   0.06462  -0.0610   0.8500   0.0170
  -5.250  -0.0106   0.04524   0.04283  -0.0576   0.7764   0.0175
  -5.000   0.0056   0.04221   0.03972  -0.0590   0.7593   0.0179
  -4.750   0.0257   0.03914   0.03656  -0.0618   0.7425   0.0186
  -4.500   0.0497   0.03582   0.03314  -0.0658   0.7263   0.0195
  -4.250   0.0888   0.03196   0.02913  -0.0742   0.7114   0.0222
  -4.000   0.1286   0.02838   0.02537  -0.0814   0.6965   0.0227
  -3.750   0.1629   0.02468   0.02150  -0.0865   0.6823   0.0228
  -3.500   0.1875   0.01944   0.01614  -0.0907   0.6686   0.0237
  -3.250   0.2138   0.01738   0.01394  -0.0927   0.6523   0.0245
  -3.000   0.2436   0.01537   0.01178  -0.0952   0.6364   0.0260
  -2.750   0.2873   0.01336   0.00946  -0.0995   0.6219   0.0306
  -2.500   0.3256   0.01080   0.00655  -0.1030   0.6080   0.0310
  -2.250   0.3571   0.00768   0.00316  -0.1068   0.5945   0.0337
  -1.000   0.5335   0.01322   0.00655  -0.1174   0.5306   0.0398
  -0.750   0.5625   0.01294   0.00604  -0.1174   0.5135   0.0416
  -0.500   0.5919   0.01229   0.00515  -0.1175   0.4972   0.0418
  -0.250   0.6206   0.01191   0.00459  -0.1176   0.4811   0.0421
   0.000   0.6496   0.01121   0.00370  -0.1178   0.4652   0.0435
   0.250   0.6782   0.01087   0.00324  -0.1178   0.4495   0.0447
   0.500   0.7065   0.01071   0.00299  -0.1179   0.4343   0.0463
   1.000   0.7629   0.01065   0.00275  -0.1179   0.4065   0.0499
   1.250   0.7910   0.01068   0.00271  -0.1179   0.3938   0.0527
   1.500   0.8190   0.01072   0.00268  -0.1179   0.3819   0.0574
   1.750   0.8473   0.01067   0.00274  -0.1180   0.3708   0.1209
   2.000   0.8755   0.01051   0.00295  -0.1183   0.3605   0.3209
   2.250   0.8975   0.00943   0.00309  -0.1173   0.3517   1.0000
   2.500   0.9250   0.00963   0.00319  -0.1172   0.3437   1.0000
   2.750   0.9524   0.00984   0.00331  -0.1171   0.3361   1.0000
   3.000   0.9798   0.01003   0.00344  -0.1170   0.3296   1.0000
   3.250   1.0071   0.01024   0.00358  -0.1169   0.3232   1.0000
   3.500   1.0342   0.01047   0.00374  -0.1168   0.3174   1.0000
   3.750   1.0615   0.01065   0.00390  -0.1167   0.3118   1.0000
   4.000   1.0882   0.01091   0.00409  -0.1165   0.3067   1.0000
   4.250   1.1153   0.01112   0.00429  -0.1164   0.3024   1.0000
   4.500   1.1423   0.01131   0.00448  -0.1163   0.2981   1.0000
   4.750   1.1688   0.01158   0.00470  -0.1162   0.2928   1.0000
   5.000   1.1958   0.01174   0.00487  -0.1161   0.2858   1.0000
   5.250   1.2221   0.01198   0.00506  -0.1159   0.2786   1.0000
   5.500   1.2488   0.01217   0.00526  -0.1158   0.2728   1.0000
   5.750   1.2752   0.01238   0.00548  -0.1156   0.2663   1.0000
   6.000   1.3016   0.01258   0.00567  -0.1155   0.2582   1.0000
   6.250   1.3276   0.01282   0.00586  -0.1153   0.2467   1.0000
   6.500   1.3536   0.01305   0.00605  -0.1152   0.2307   1.0000
   6.750   1.3769   0.01365   0.00638  -0.1147   0.1902   1.0000
   7.000   1.3803   0.01705   0.00875  -0.1120   0.0236   1.0000
   7.250   1.4031   0.01768   0.00941  -0.1113   0.0188   1.0000
   7.500   1.4260   0.01826   0.01013  -0.1106   0.0166   1.0000
   7.750   1.4481   0.01889   0.01088  -0.1098   0.0156   1.0000
   8.000   1.4689   0.01966   0.01178  -0.1088   0.0145   1.0000
   8.250   1.4881   0.02057   0.01282  -0.1077   0.0137   1.0000
   8.500   1.5051   0.02165   0.01403  -0.1062   0.0130   1.0000
   8.750   1.5191   0.02295   0.01547  -0.1045   0.0125   1.0000
   9.000   1.5287   0.02451   0.01718  -0.1022   0.0121   1.0000
   9.250   1.5310   0.02646   0.01931  -0.0990   0.0117   1.0000
   9.500   1.5210   0.02874   0.02175  -0.0943   0.0114   1.0000
   9.750   1.5191   0.03063   0.02377  -0.0913   0.0110   1.0000
  10.000   1.5253   0.03214   0.02538  -0.0895   0.0107   1.0000
  10.250   1.5234   0.03454   0.02790  -0.0876   0.0106   1.0000
  10.500   1.5201   0.03739   0.03088  -0.0863   0.0105   1.0000
  10.750   1.5160   0.04063   0.03424  -0.0856   0.0104   1.0000
  11.000   1.5110   0.04423   0.03796  -0.0853   0.0103   1.0000
  11.250   1.5055   0.04804   0.04188  -0.0852   0.0102   1.0000
  11.500   1.4998   0.05195   0.04590  -0.0852   0.0102   1.0000
  11.750   1.4945   0.05580   0.04984  -0.0850   0.0101   1.0000
  12.000   1.4897   0.05948   0.05361  -0.0846   0.0101   1.0000
  12.250   1.4865   0.06283   0.05703  -0.0839   0.0102   1.0000
  12.500   1.4858   0.06558   0.05985  -0.0824   0.0103   1.0000
  13.000   1.2393   0.08370   0.07882  -0.0716   0.0106   1.0000
  13.250   1.2352   0.08718   0.08236  -0.0710   0.0106   1.0000
  13.500   1.2344   0.09002   0.08525  -0.0701   0.0105   1.0000
  13.750   1.2390   0.09163   0.08689  -0.0684   0.0105   1.0000
  14.000   1.2698   0.08859   0.08375  -0.0633   0.0119   1.0000
  14.750   1.3000   0.09104   0.08667  -0.0547   0.0156   1.0000
  15.000   1.2817   0.09430   0.09007  -0.0545   0.0155   1.0000
  15.250   1.2680   0.09772   0.09363  -0.0549   0.0152   1.0000
<< Back to 20-32C AIRFOIL (2032c-il)

Polar data table (+)

Polar graphs


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