20-32C AIRFOIL (2032c-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: 20-32C AIRFOIL (2032c-il) Reynolds number: 50,000 Max Cl/Cd: 40.82 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-2032c-il-50000-n5.txt Download as CSV file: xf-2032c-il-50000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-6.750 -0.2164 0.10214 0.09678 -0.0202 1.0000 0.0621
-6.500 -0.2233 0.10113 0.09590 -0.0187 1.0000 0.0634
-6.250 -0.2313 0.10033 0.09523 -0.0174 1.0000 0.0647
-6.000 -0.2134 0.09885 0.09379 -0.0250 0.9918 0.0671
-5.750 -0.1645 0.09639 0.09127 -0.0423 0.9769 0.0683
-5.500 -0.1546 0.08928 0.08418 -0.0375 0.9697 0.0706
-5.250 -0.1264 0.08495 0.07982 -0.0419 0.9586 0.0749
-5.000 -0.0770 0.08210 0.07686 -0.0565 0.9440 0.0821
-4.750 -0.0477 0.07701 0.07178 -0.0609 0.9335 0.0849
-4.500 -0.0121 0.07301 0.06772 -0.0666 0.9229 0.0949
-4.250 0.0259 0.06883 0.06349 -0.0739 0.9107 0.1009
-4.000 0.0869 0.06619 0.06057 -0.0884 0.8964 0.1125
-3.750 0.1013 0.06143 0.05587 -0.0864 0.8834 0.1157
-3.500 0.1521 0.05823 0.05245 -0.0961 0.8693 0.1279
-3.250 0.1846 0.05478 0.04891 -0.0994 0.8551 0.1431
-2.750 0.2824 0.04530 0.03876 -0.1124 0.8270 0.0778
-2.500 0.3292 0.04143 0.03452 -0.1180 0.8120 0.0727
-2.250 0.3704 0.03845 0.03118 -0.1217 0.7963 0.0753
-2.000 0.4153 0.03506 0.02725 -0.1257 0.7807 0.0723
-1.750 0.4519 0.03278 0.02450 -0.1277 0.7631 0.0752
-1.500 0.4893 0.03067 0.02180 -0.1295 0.7456 0.0799
-1.250 0.5242 0.02876 0.01929 -0.1305 0.7283 0.0802
-1.000 0.5569 0.02724 0.01719 -0.1309 0.7111 0.0813
-0.750 0.5880 0.02606 0.01544 -0.1308 0.6941 0.0832
-0.500 0.6156 0.02539 0.01456 -0.1305 0.6765 0.0920
-0.250 0.6444 0.02478 0.01337 -0.1298 0.6585 0.0976
0.000 0.6717 0.02418 0.01250 -0.1291 0.6411 0.1014
0.250 0.6988 0.02376 0.01177 -0.1284 0.6242 0.1068
0.500 0.7259 0.02342 0.01117 -0.1277 0.6077 0.1146
0.750 0.7530 0.02317 0.01076 -0.1271 0.5916 0.1297
1.000 0.7804 0.02287 0.01053 -0.1268 0.5760 0.1799
1.250 0.8016 0.02135 0.01047 -0.1255 0.5614 1.0000
1.500 0.8283 0.02171 0.01035 -0.1249 0.5464 1.0000
1.750 0.8545 0.02210 0.01037 -0.1244 0.5319 1.0000
2.000 0.8806 0.02251 0.01049 -0.1239 0.5180 1.0000
2.250 0.9065 0.02295 0.01068 -0.1235 0.5048 1.0000
2.500 0.9323 0.02341 0.01094 -0.1231 0.4922 1.0000
2.750 0.9580 0.02388 0.01121 -0.1226 0.4806 1.0000
3.000 0.9839 0.02435 0.01147 -0.1222 0.4703 1.0000
3.250 1.0092 0.02489 0.01192 -0.1218 0.4595 1.0000
3.500 1.0344 0.02545 0.01239 -0.1215 0.4495 1.0000
3.750 1.0601 0.02597 0.01276 -0.1210 0.4410 1.0000
4.000 1.0848 0.02660 0.01341 -0.1207 0.4316 1.0000
4.250 1.1100 0.02721 0.01396 -0.1203 0.4239 1.0000
4.500 1.1347 0.02785 0.01459 -0.1199 0.4161 1.0000
4.750 1.1594 0.02853 0.01526 -0.1195 0.4089 1.0000
5.000 1.1837 0.02922 0.01602 -0.1191 0.4016 1.0000
5.250 1.2083 0.02992 0.01673 -0.1188 0.3955 1.0000
5.500 1.2316 0.03076 0.01768 -0.1183 0.3888 1.0000
5.750 1.2568 0.03142 0.01838 -0.1180 0.3837 1.0000
6.000 1.2786 0.03242 0.01959 -0.1174 0.3773 1.0000
6.250 1.3021 0.03324 0.02050 -0.1170 0.3718 1.0000
6.500 1.3259 0.03406 0.02141 -0.1165 0.3670 1.0000
6.750 1.3460 0.03525 0.02293 -0.1159 0.3614 1.0000
7.000 1.3689 0.03619 0.02402 -0.1154 0.3569 1.0000
7.250 1.3927 0.03708 0.02503 -0.1149 0.3530 1.0000
7.500 1.4091 0.03859 0.02692 -0.1140 0.3475 1.0000
7.750 1.4326 0.03866 0.02699 -0.1130 0.3385 1.0000
8.000 1.4448 0.03873 0.02724 -0.1108 0.3218 1.0000
8.250 1.4590 0.03893 0.02754 -0.1089 0.3081 1.0000
8.500 1.4721 0.03951 0.02830 -0.1070 0.2967 1.0000
8.750 1.4805 0.04030 0.02932 -0.1048 0.2838 1.0000
9.000 1.4838 0.04093 0.03002 -0.1019 0.2668 1.0000
9.250 1.4869 0.04226 0.03161 -0.0995 0.2542 1.0000
9.500 1.4862 0.04379 0.03337 -0.0968 0.2416 1.0000
9.750 1.4817 0.04569 0.03546 -0.0942 0.2269 1.0000
10.000 1.4769 0.04808 0.03804 -0.0923 0.2115 1.0000
10.500 1.4631 0.05460 0.04500 -0.0906 0.1709 1.0000
11.250 1.4032 0.07162 0.06159 -0.0927 0.0644 1.0000
11.500 1.3826 0.07833 0.06839 -0.0944 0.0560 1.0000
11.750 1.3627 0.08518 0.07537 -0.0962 0.0520 1.0000
12.000 1.3429 0.09221 0.08254 -0.0982 0.0483 1.0000
|
Polar data table (+)
Polar graphs
<< Back to 20-32C AIRFOIL (2032c-il)