EPPLER 908 AIRFOIL (e908-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 908 AIRFOIL (e908-il) Reynolds number: 500,000 Max Cl/Cd: 91.8 at α=2° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e908-il-500000-n5.txt Download as CSV file: xf-e908-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 908 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3114 0.06896 0.06657 -0.0904 0.9690 0.0022
-8.000 -0.3160 0.06494 0.06258 -0.0929 0.9631 0.0022
-7.750 -0.3117 0.05714 0.05478 -0.1042 0.9592 0.0021
-7.500 -0.3095 0.05111 0.04864 -0.1088 0.9549 0.0021
-7.250 -0.3070 0.04588 0.04325 -0.1108 0.9504 0.0020
-7.000 -0.2957 0.04014 0.03727 -0.1134 0.9481 0.0020
-6.750 -0.2804 0.03427 0.03104 -0.1156 0.9465 0.0019
-6.500 -0.2612 0.02924 0.02553 -0.1170 0.9454 0.0018
-6.250 -0.2599 0.02662 0.02262 -0.1128 0.9394 0.0018
-6.000 -0.2409 0.02302 0.01856 -0.1121 0.9373 0.0017
-5.750 -0.2174 0.01981 0.01487 -0.1118 0.9360 0.0017
-5.500 -0.1910 0.01734 0.01198 -0.1118 0.9351 0.0017
-5.250 -0.1633 0.01533 0.00964 -0.1119 0.9343 0.0019
-5.000 -0.1346 0.01391 0.00800 -0.1122 0.9337 0.0021
-4.750 -0.1049 0.01311 0.00706 -0.1129 0.9331 0.0025
-4.250 -0.0663 0.01184 0.00564 -0.1099 0.9272 0.0042
-4.000 -0.0385 0.01167 0.00547 -0.1103 0.9256 0.0066
-3.750 -0.0097 0.01099 0.00465 -0.1106 0.9242 0.0071
-3.500 0.0204 0.01026 0.00368 -0.1111 0.9230 0.0149
-3.250 0.0509 0.00964 0.00324 -0.1121 0.9219 0.0768
-3.000 0.0827 0.00911 0.00298 -0.1135 0.9209 0.1607
-2.750 0.1152 0.00872 0.00276 -0.1150 0.9200 0.2219
-2.500 0.1340 0.00831 0.00269 -0.1136 0.9162 0.3261
-2.250 0.1575 0.00779 0.00259 -0.1132 0.9133 0.4557
-2.000 0.1853 0.00743 0.00246 -0.1135 0.9113 0.5534
-1.750 0.2158 0.00727 0.00242 -0.1143 0.9098 0.6110
-1.500 0.2475 0.00719 0.00235 -0.1153 0.9086 0.6377
-1.250 0.2797 0.00711 0.00229 -0.1164 0.9074 0.6580
-1.000 0.3050 0.00711 0.00230 -0.1160 0.9047 0.6715
-0.750 0.3277 0.00712 0.00234 -0.1150 0.9011 0.6822
-0.500 0.3561 0.00709 0.00233 -0.1152 0.8987 0.6941
-0.250 0.3866 0.00703 0.00231 -0.1159 0.8968 0.7098
0.000 0.4181 0.00696 0.00230 -0.1169 0.8951 0.7295
0.250 0.4470 0.00691 0.00233 -0.1172 0.8929 0.7462
0.500 0.4668 0.00693 0.00244 -0.1155 0.8881 0.7594
0.750 0.4952 0.00690 0.00246 -0.1157 0.8850 0.7703
1.000 0.5262 0.00685 0.00251 -0.1165 0.8826 0.7782
1.250 0.5554 0.00682 0.00254 -0.1170 0.8798 0.7854
1.500 0.5765 0.00683 0.00264 -0.1155 0.8741 0.7929
1.750 0.6069 0.00669 0.00256 -0.1160 0.8662 0.7995
2.000 0.6215 0.00677 0.00185 -0.1119 0.6896 0.8075
2.250 0.6025 0.00835 0.00232 -0.1020 0.4562 0.8196
2.500 0.5864 0.01129 0.00333 -0.0941 0.0064 0.8334
2.750 0.6092 0.01148 0.00359 -0.0931 0.0027 0.8465
3.250 0.6507 0.01194 0.00446 -0.0899 0.0024 0.8862
3.500 0.6681 0.01228 0.00502 -0.0875 0.0022 0.9288
3.750 0.6953 0.01294 0.00583 -0.0875 0.0022 1.0000
4.000 0.7133 0.01380 0.00679 -0.0854 0.0022 1.0000
4.250 0.7319 0.01477 0.00784 -0.0835 0.0022 1.0000
4.500 0.7538 0.01607 0.00924 -0.0822 0.0022 1.0000
4.750 0.7825 0.01773 0.01101 -0.0822 0.0023 1.0000
5.000 0.8166 0.01985 0.01331 -0.0831 0.0025 1.0000
5.250 0.8504 0.02258 0.01629 -0.0837 0.0027 1.0000
7.250 0.9552 0.03775 0.03416 -0.0595 0.0047 1.0000
7.500 0.9576 0.04155 0.03818 -0.0560 0.0046 1.0000
7.750 0.9568 0.04552 0.04237 -0.0523 0.0045 1.0000
8.000 0.9529 0.04948 0.04651 -0.0485 0.0044 1.0000
8.250 0.9455 0.05330 0.05050 -0.0446 0.0043 1.0000
8.500 0.9318 0.05668 0.05403 -0.0399 0.0043 1.0000
8.750 0.9148 0.06002 0.05750 -0.0355 0.0043 1.0000
9.000 0.8950 0.06379 0.06140 -0.0318 0.0043 1.0000
9.250 0.8740 0.06789 0.06563 -0.0291 0.0043 1.0000
9.500 0.8503 0.07280 0.07067 -0.0274 0.0044 1.0000
9.750 0.8258 0.07829 0.07627 -0.0273 0.0044 1.0000
10.000 0.7998 0.08490 0.08298 -0.0290 0.0045 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 908 AIRFOIL (e908-il)