EPPLER 398 AIRFOIL (e398-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 398 AIRFOIL (e398-il) Reynolds number: 100,000 Max Cl/Cd: 47.52 at α=10.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e398-il-100000.txt Download as CSV file: xf-e398-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 398 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.2291 0.10400 0.09975 -0.0640 0.9660 0.1303
-8.500 -0.2999 0.08289 0.07876 -0.0827 0.9528 0.0704
-8.250 -0.3657 0.06363 0.05934 -0.1024 0.9353 0.0598
-8.000 -0.3674 0.05601 0.05143 -0.1104 0.9263 0.0592
-7.750 -0.3658 0.04881 0.04364 -0.1169 0.9170 0.0590
-7.500 -0.3553 0.04293 0.03690 -0.1213 0.9088 0.0602
-7.250 -0.3278 0.04180 0.03599 -0.1224 0.9031 0.0662
-7.000 -0.2919 0.03711 0.03053 -0.1279 0.8996 0.0735
-6.750 -0.2744 0.03601 0.02917 -0.1274 0.8910 0.0809
-6.500 -0.2397 0.03416 0.02713 -0.1300 0.8864 0.0916
-6.250 -0.1973 0.03235 0.02503 -0.1336 0.8837 0.1054
-6.000 -0.1825 0.03157 0.02411 -0.1323 0.8752 0.1143
-5.750 -0.1461 0.03041 0.02277 -0.1344 0.8708 0.1287
-5.500 -0.1021 0.02927 0.02149 -0.1377 0.8682 0.1465
-5.250 -0.0875 0.02897 0.02110 -0.1360 0.8598 0.1584
-5.000 -0.0527 0.02834 0.02041 -0.1376 0.8554 0.1775
-4.750 -0.0086 0.02752 0.01946 -0.1407 0.8529 0.2022
-4.500 0.0068 0.02762 0.01956 -0.1391 0.8449 0.2186
-4.250 0.0389 0.02734 0.01930 -0.1401 0.8401 0.2415
-4.000 0.0812 0.02684 0.01879 -0.1427 0.8375 0.2694
-3.750 0.1263 0.02635 0.01825 -0.1457 0.8356 0.2996
-3.500 0.1262 0.02708 0.01895 -0.1416 0.8248 0.3137
-3.250 0.1671 0.02674 0.01863 -0.1438 0.8220 0.3424
-3.000 0.2096 0.02633 0.01825 -0.1461 0.8199 0.3712
-2.750 0.2072 0.02730 0.01924 -0.1419 0.8095 0.3846
-2.500 0.2465 0.02703 0.01894 -0.1437 0.8064 0.4138
-2.250 0.2891 0.02664 0.01852 -0.1459 0.8043 0.4443
-2.000 0.2842 0.02783 0.01980 -0.1415 0.7941 0.4569
-1.750 0.3208 0.02764 0.01965 -0.1427 0.7908 0.4848
-1.500 0.3623 0.02725 0.01927 -0.1446 0.7886 0.5151
-1.250 0.3552 0.02867 0.02072 -0.1400 0.7783 0.5297
-1.000 0.3913 0.02853 0.02061 -0.1412 0.7751 0.5603
-0.750 0.4315 0.02815 0.02028 -0.1426 0.7729 0.5920
-0.500 0.4189 0.02984 0.02205 -0.1374 0.7621 0.6074
-0.250 0.4546 0.02966 0.02194 -0.1382 0.7591 0.6410
0.000 0.4941 0.02925 0.02160 -0.1394 0.7570 0.6774
0.250 0.4742 0.03130 0.02377 -0.1333 0.7455 0.6962
0.500 0.5086 0.03101 0.02355 -0.1335 0.7428 0.7385
0.750 0.5452 0.03047 0.02311 -0.1337 0.7409 0.7858
1.000 0.5162 0.03287 0.02566 -0.1264 0.7287 0.8196
1.250 0.5381 0.03221 0.02517 -0.1237 0.7261 0.9070
1.500 0.5989 0.03152 0.02433 -0.1294 0.7246 1.0000
1.750 0.5828 0.03445 0.02718 -0.1259 0.7120 1.0000
2.000 0.6304 0.03409 0.02668 -0.1288 0.7100 1.0000
2.250 0.6809 0.03352 0.02600 -0.1317 0.7086 1.0000
2.500 0.6569 0.03681 0.02926 -0.1270 0.6954 1.0000
2.750 0.6768 0.03785 0.03024 -0.1266 0.6895 1.0000
3.000 0.6859 0.03938 0.03175 -0.1253 0.6812 1.0000
3.250 0.7288 0.03905 0.03136 -0.1269 0.6787 1.0000
3.500 0.7774 0.03831 0.03058 -0.1291 0.6772 1.0000
3.750 0.7556 0.04173 0.03401 -0.1249 0.6640 1.0000
4.000 0.7559 0.04391 0.03619 -0.1229 0.6544 1.0000
4.250 0.7855 0.04426 0.03653 -0.1232 0.6493 1.0000
4.500 0.8298 0.04365 0.03592 -0.1245 0.6469 1.0000
4.750 0.8762 0.04284 0.03513 -0.1260 0.6450 1.0000
5.000 0.8599 0.04609 0.03840 -0.1227 0.6318 1.0000
5.250 0.8529 0.04884 0.04117 -0.1202 0.6200 1.0000
5.500 0.8929 0.04832 0.04068 -0.1210 0.6166 1.0000
5.750 0.9424 0.04705 0.03947 -0.1223 0.6149 1.0000
6.000 0.9279 0.05038 0.04283 -0.1194 0.6014 1.0000
6.250 0.9291 0.05264 0.04512 -0.1177 0.5904 1.0000
6.500 0.9666 0.05205 0.04459 -0.1180 0.5861 1.0000
6.750 1.0166 0.05036 0.04299 -0.1189 0.5841 1.0000
7.000 1.0715 0.04807 0.04080 -0.1200 0.5831 1.0000
7.250 1.1328 0.04511 0.03797 -0.1215 0.5826 1.0000
7.500 1.1045 0.04962 0.04252 -0.1176 0.5659 1.0000
7.750 1.1719 0.04581 0.03886 -0.1192 0.5663 1.0000
8.000 1.2561 0.04123 0.03444 -0.1229 0.5668 1.0000
8.250 1.2383 0.04397 0.03724 -0.1184 0.5530 1.0000
8.500 1.3386 0.03834 0.03181 -0.1239 0.5525 1.0000
8.750 1.3178 0.04080 0.03435 -0.1184 0.5390 1.0000
9.000 1.4557 0.03357 0.02730 -0.1291 0.5361 1.0000
9.250 1.4410 0.03491 0.02876 -0.1231 0.5234 1.0000
9.500 1.4526 0.03491 0.02887 -0.1202 0.5113 1.0000
9.750 1.4881 0.03350 0.02758 -0.1199 0.4988 1.0000
10.000 1.5152 0.03253 0.02669 -0.1187 0.4846 1.0000
10.250 1.5302 0.03222 0.02645 -0.1162 0.4691 1.0000
10.500 1.5379 0.03236 0.02666 -0.1129 0.4523 1.0000
10.750 1.5444 0.03266 0.02700 -0.1097 0.4338 1.0000
11.000 1.5543 0.03286 0.02716 -0.1070 0.4130 1.0000
11.250 1.5457 0.03446 0.02882 -0.1030 0.3924 1.0000
11.500 1.5453 0.03568 0.02997 -0.0998 0.3690 1.0000
11.750 1.5374 0.03772 0.03198 -0.0966 0.3453 1.0000
12.000 1.5321 0.03974 0.03387 -0.0938 0.3206 1.0000
12.250 1.5223 0.04237 0.03643 -0.0911 0.2961 1.0000
12.500 1.5130 0.04518 0.03912 -0.0887 0.2721 1.0000
12.750 1.5030 0.04824 0.04204 -0.0866 0.2490 1.0000
13.000 1.4927 0.05157 0.04526 -0.0849 0.2269 1.0000
13.250 1.4827 0.05505 0.04865 -0.0834 0.2061 1.0000
13.500 1.4738 0.05859 0.05201 -0.0821 0.1868 1.0000
13.750 1.4651 0.06235 0.05573 -0.0812 0.1682 1.0000
14.000 1.4575 0.06611 0.05941 -0.0804 0.1513 1.0000
14.250 1.4512 0.06985 0.06306 -0.0798 0.1360 1.0000
14.500 1.4464 0.07353 0.06667 -0.0793 0.1221 1.0000
14.750 1.4435 0.07707 0.07015 -0.0788 0.1095 1.0000
15.000 1.4418 0.08053 0.07358 -0.0784 0.0985 1.0000
15.250 1.4430 0.08369 0.07669 -0.0780 0.0888 1.0000
15.500 1.4475 0.08638 0.07922 -0.0775 0.0798 1.0000
15.750 1.4453 0.09022 0.08332 -0.0777 0.0734 1.0000
16.000 1.4512 0.09288 0.08592 -0.0773 0.0668 1.0000
16.250 1.4518 0.09635 0.08953 -0.0775 0.0617 1.0000
16.500 1.4596 0.09891 0.09208 -0.0771 0.0565 1.0000
16.750 1.4585 0.10273 0.09613 -0.0776 0.0529 1.0000
17.000 1.4722 0.10455 0.09781 -0.0770 0.0484 1.0000
17.250 1.4643 0.10939 0.10301 -0.0782 0.0464 1.0000
17.500 1.4608 0.11366 0.10749 -0.0792 0.0441 1.0000
17.750 1.4696 0.11609 0.10990 -0.0793 0.0416 1.0000
18.000 1.4668 0.12064 0.11467 -0.0804 0.0401 1.0000
18.250 1.4525 0.12671 0.12109 -0.0829 0.0394 1.0000
18.500 1.4364 0.13325 0.12794 -0.0860 0.0388 1.0000
18.750 1.4199 0.14015 0.13513 -0.0897 0.0383 1.0000
19.000 1.4014 0.14773 0.14298 -0.0940 0.0381 1.0000
19.250 1.3799 0.15627 0.15178 -0.0994 0.0382 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 398 AIRFOIL (e398-il)