EPPLER 472 AIRFOIL (e472-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 472 AIRFOIL (e472-il) Reynolds number: 100,000 Max Cl/Cd: 33.6 at α=8° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e472-il-100000-n5.txt Download as CSV file: xf-e472-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 472 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.250 -1.0040 0.06044 0.05406 -0.0384 1.0000 0.0942
-13.000 -1.0640 0.05086 0.04414 -0.0436 1.0000 0.0945
-12.750 -1.0947 0.04643 0.03941 -0.0427 1.0000 0.0956
-12.500 -1.1191 0.04323 0.03586 -0.0395 1.0000 0.0971
-12.250 -1.1068 0.04270 0.03544 -0.0377 1.0000 0.0989
-12.000 -1.0984 0.04198 0.03472 -0.0355 1.0000 0.1010
-11.750 -1.0941 0.04069 0.03331 -0.0333 1.0000 0.1032
-11.500 -1.0927 0.03896 0.03133 -0.0308 1.0000 0.1059
-11.250 -1.0934 0.03686 0.02883 -0.0281 1.0000 0.1088
-11.000 -1.0780 0.03613 0.02813 -0.0265 1.0000 0.1111
-10.750 -1.0621 0.03552 0.02751 -0.0249 1.0000 0.1136
-10.500 -1.0489 0.03456 0.02642 -0.0231 1.0000 0.1165
-10.250 -1.0381 0.03321 0.02480 -0.0210 1.0000 0.1199
-10.000 -1.0247 0.03204 0.02341 -0.0192 1.0000 0.1230
-9.750 -1.0057 0.03152 0.02292 -0.0178 1.0000 0.1256
-9.500 -0.9881 0.03084 0.02217 -0.0163 1.0000 0.1286
-9.250 -0.9719 0.02989 0.02102 -0.0146 1.0000 0.1323
-9.000 -0.9560 0.02883 0.01968 -0.0128 1.0000 0.1361
-8.750 -0.9356 0.02831 0.01921 -0.0116 1.0000 0.1388
-8.500 -0.9159 0.02772 0.01858 -0.0102 1.0000 0.1420
-8.250 -0.8969 0.02699 0.01769 -0.0087 1.0000 0.1459
-8.000 -0.8782 0.02614 0.01658 -0.0072 1.0000 0.1501
-7.750 -0.8569 0.02560 0.01611 -0.0060 1.0000 0.1530
-7.500 -0.8360 0.02507 0.01554 -0.0047 1.0000 0.1565
-7.250 -0.8153 0.02445 0.01480 -0.0034 1.0000 0.1607
-7.000 -0.7946 0.02379 0.01393 -0.0021 1.0000 0.1650
-6.750 -0.7728 0.02327 0.01348 -0.0009 1.0000 0.1683
-6.500 -0.7510 0.02277 0.01296 0.0003 1.0000 0.1721
-6.250 -0.7292 0.02224 0.01233 0.0015 1.0000 0.1766
-6.000 -0.7073 0.02171 0.01168 0.0027 1.0000 0.1809
-5.750 -0.6851 0.02124 0.01128 0.0039 1.0000 0.1847
-5.500 -0.6628 0.02079 0.01081 0.0051 1.0000 0.1889
-5.250 -0.6403 0.02035 0.01027 0.0062 1.0000 0.1938
-5.000 -0.6179 0.01989 0.00981 0.0074 1.0000 0.1980
-4.750 -0.5956 0.01949 0.00947 0.0085 1.0000 0.2025
-4.500 -0.5730 0.01911 0.00906 0.0097 1.0000 0.2075
-4.250 -0.5502 0.01874 0.00863 0.0108 1.0000 0.2125
-4.000 -0.5280 0.01835 0.00835 0.0120 1.0000 0.2171
-3.750 -0.5056 0.01803 0.00805 0.0132 1.0000 0.2228
-3.500 -0.4829 0.01773 0.00772 0.0143 1.0000 0.2285
-3.250 -0.4609 0.01738 0.00749 0.0155 1.0000 0.2336
-3.000 -0.4388 0.01710 0.00725 0.0167 1.0000 0.2399
-2.750 -0.4167 0.01684 0.00703 0.0179 1.0000 0.2465
-2.500 -0.3952 0.01658 0.00688 0.0192 1.0000 0.2530
-2.250 -0.3554 0.01635 0.00671 0.0168 0.9914 0.2624
-2.000 -0.3103 0.01603 0.00653 0.0134 0.9739 0.2736
-1.750 -0.2726 0.01566 0.00630 0.0116 0.9501 0.2852
-1.500 -0.2350 0.01529 0.00605 0.0101 0.9210 0.2987
-1.250 -0.1885 0.01491 0.00580 0.0069 0.8831 0.3163
-1.000 -0.1276 0.01452 0.00549 0.0009 0.8158 0.3432
-0.750 -0.0782 0.01436 0.00518 -0.0025 0.7158 0.3740
-0.500 -0.0490 0.01442 0.00504 -0.0021 0.6266 0.4057
-0.250 -0.0243 0.01448 0.00497 -0.0011 0.5531 0.4447
0.000 0.0000 0.01450 0.00495 0.0000 0.4932 0.4933
0.250 0.0243 0.01449 0.00497 0.0010 0.4446 0.5530
0.500 0.0491 0.01441 0.00504 0.0021 0.4058 0.6268
0.750 0.0782 0.01436 0.00518 0.0025 0.3738 0.7160
1.000 0.1278 0.01452 0.00549 -0.0010 0.3432 0.8162
1.250 0.1885 0.01491 0.00580 -0.0069 0.3163 0.8831
1.500 0.2350 0.01529 0.00605 -0.0101 0.2987 0.9209
1.750 0.2726 0.01566 0.00630 -0.0116 0.2853 0.9500
2.000 0.3104 0.01603 0.00653 -0.0134 0.2736 0.9740
2.250 0.3557 0.01635 0.00671 -0.0168 0.2624 0.9916
2.500 0.3952 0.01658 0.00688 -0.0192 0.2530 1.0000
2.750 0.4166 0.01684 0.00703 -0.0179 0.2464 1.0000
3.000 0.4387 0.01710 0.00725 -0.0167 0.2398 1.0000
3.250 0.4609 0.01737 0.00748 -0.0155 0.2335 1.0000
3.500 0.4828 0.01772 0.00771 -0.0143 0.2286 1.0000
3.750 0.5056 0.01803 0.00805 -0.0132 0.2228 1.0000
4.000 0.5280 0.01835 0.00835 -0.0120 0.2171 1.0000
4.250 0.5502 0.01873 0.00863 -0.0108 0.2125 1.0000
4.500 0.5730 0.01911 0.00906 -0.0097 0.2075 1.0000
4.750 0.5956 0.01949 0.00947 -0.0085 0.2025 1.0000
5.000 0.6179 0.01989 0.00981 -0.0074 0.1980 1.0000
5.250 0.6403 0.02035 0.01027 -0.0062 0.1938 1.0000
5.500 0.6628 0.02079 0.01081 -0.0051 0.1889 1.0000
5.750 0.6851 0.02124 0.01128 -0.0039 0.1847 1.0000
6.000 0.7073 0.02171 0.01168 -0.0028 0.1810 1.0000
6.250 0.7292 0.02225 0.01233 -0.0015 0.1766 1.0000
6.500 0.7510 0.02277 0.01297 -0.0003 0.1721 1.0000
6.750 0.7728 0.02327 0.01348 0.0009 0.1683 1.0000
7.000 0.7947 0.02379 0.01393 0.0021 0.1650 1.0000
7.250 0.8153 0.02445 0.01480 0.0034 0.1607 1.0000
7.500 0.8360 0.02506 0.01554 0.0047 0.1565 1.0000
7.750 0.8569 0.02560 0.01611 0.0060 0.1530 1.0000
8.000 0.8783 0.02614 0.01658 0.0072 0.1501 1.0000
8.250 0.8969 0.02699 0.01769 0.0087 0.1460 1.0000
8.500 0.9159 0.02772 0.01858 0.0102 0.1420 1.0000
8.750 0.9356 0.02831 0.01921 0.0116 0.1388 1.0000
9.000 0.9560 0.02884 0.01968 0.0128 0.1361 1.0000
9.250 0.9720 0.02990 0.02102 0.0146 0.1324 1.0000
9.500 0.9881 0.03083 0.02216 0.0163 0.1286 1.0000
9.750 1.0057 0.03153 0.02293 0.0178 0.1256 1.0000
10.000 1.0248 0.03203 0.02339 0.0192 0.1230 1.0000
10.250 1.0381 0.03322 0.02480 0.0210 0.1199 1.0000
10.500 1.0490 0.03456 0.02642 0.0231 0.1165 1.0000
10.750 1.0622 0.03552 0.02751 0.0249 0.1136 1.0000
11.000 1.0782 0.03612 0.02812 0.0265 0.1111 1.0000
11.250 1.0934 0.03687 0.02884 0.0281 0.1088 1.0000
11.500 1.0929 0.03896 0.03133 0.0308 0.1059 1.0000
11.750 1.0943 0.04069 0.03331 0.0333 0.1032 1.0000
12.000 1.0987 0.04198 0.03472 0.0355 0.1010 1.0000
12.250 1.1073 0.04269 0.03542 0.0376 0.0989 1.0000
12.500 1.1190 0.04325 0.03589 0.0395 0.0970 1.0000
12.750 1.0949 0.04644 0.03942 0.0427 0.0956 1.0000
13.000 1.0648 0.05085 0.04413 0.0435 0.0945 1.0000
13.250 1.0046 0.06045 0.05407 0.0383 0.0941 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 472 AIRFOIL (e472-il)