EPPLER 1210 AIRFOIL (e1210-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 1210 AIRFOIL (e1210-il) Reynolds number: 100,000 Max Cl/Cd: 45.9 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e1210-il-100000-n5.txt Download as CSV file: xf-e1210-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 1210 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.2313 0.10979 0.10473 -0.0362 1.0000 0.0475
-10.500 -0.2352 0.10550 0.10049 -0.0377 1.0000 0.0482
-10.250 -0.2418 0.10090 0.09596 -0.0394 1.0000 0.0497
-10.000 -0.2346 0.09918 0.09432 -0.0392 1.0000 0.0509
-9.750 -0.2364 0.09592 0.09115 -0.0400 1.0000 0.0524
-9.500 -0.2467 0.09034 0.08564 -0.0427 0.9943 0.0537
-9.250 -0.2303 0.08545 0.08077 -0.0486 0.9496 0.0557
-9.000 -0.2122 0.07910 0.07433 -0.0576 0.9108 0.0588
-8.750 -0.1989 0.07191 0.06703 -0.0674 0.8782 0.0615
-8.500 -0.3449 0.03797 0.03197 -0.0988 0.8434 0.0663
-8.250 -0.3092 0.03911 0.03322 -0.0984 0.8204 0.0680
-8.000 -0.2960 0.03652 0.03024 -0.0994 0.7998 0.0709
-7.750 -0.2835 0.03362 0.02685 -0.1001 0.7817 0.0740
-7.500 -0.2601 0.03308 0.02618 -0.0997 0.7640 0.0764
-7.250 -0.2448 0.03039 0.02288 -0.0999 0.7484 0.0805
-7.000 -0.2191 0.03024 0.02273 -0.0994 0.7331 0.0826
-6.750 -0.1963 0.02915 0.02132 -0.0991 0.7195 0.0861
-6.500 -0.1740 0.02795 0.01982 -0.0988 0.7053 0.0894
-6.250 -0.1489 0.02756 0.01934 -0.0983 0.6923 0.0922
-6.000 -0.1256 0.02636 0.01768 -0.0979 0.6802 0.0966
-5.750 -0.1005 0.02589 0.01717 -0.0974 0.6675 0.0992
-5.500 -0.0750 0.02543 0.01653 -0.0968 0.6566 0.1027
-5.250 -0.0503 0.02451 0.01524 -0.0964 0.6445 0.1071
-5.000 -0.0244 0.02420 0.01492 -0.0959 0.6340 0.1099
-4.750 0.0014 0.02375 0.01433 -0.0954 0.6227 0.1138
-4.500 0.0274 0.02313 0.01338 -0.0949 0.6131 0.1183
-4.250 0.0533 0.02280 0.01308 -0.0944 0.6021 0.1216
-4.000 0.0797 0.02244 0.01255 -0.0939 0.5930 0.1259
-3.750 0.1061 0.02200 0.01190 -0.0935 0.5828 0.1307
-3.500 0.1324 0.02170 0.01156 -0.0929 0.5743 0.1345
-3.250 0.1589 0.02141 0.01120 -0.0925 0.5642 0.1394
-3.000 0.1857 0.02110 0.01067 -0.0920 0.5561 0.1444
-2.750 0.2119 0.02083 0.01046 -0.0915 0.5468 0.1489
-2.500 0.2385 0.02062 0.01013 -0.0910 0.5384 0.1546
-2.250 0.2653 0.02039 0.00979 -0.0905 0.5303 0.1601
-2.000 0.2915 0.02019 0.00962 -0.0900 0.5218 0.1658
-1.750 0.3183 0.02004 0.00932 -0.0895 0.5148 0.1726
-1.500 0.3445 0.01986 0.00919 -0.0890 0.5062 0.1791
-1.250 0.3710 0.01974 0.00902 -0.0885 0.4989 0.1868
-1.000 0.3974 0.01964 0.00887 -0.0880 0.4918 0.1952
-0.750 0.4236 0.01954 0.00880 -0.0875 0.4840 0.2046
-0.500 0.4500 0.01946 0.00866 -0.0870 0.4776 0.2154
-0.250 0.4761 0.01941 0.00864 -0.0865 0.4703 0.2288
0.000 0.5022 0.01934 0.00863 -0.0860 0.4634 0.2445
0.250 0.5286 0.01930 0.00856 -0.0855 0.4577 0.2655
0.750 0.5802 0.01920 0.00871 -0.0845 0.4440 0.3370
1.000 0.6063 0.01908 0.00874 -0.0840 0.4387 0.4052
1.250 0.6306 0.01893 0.00896 -0.0833 0.4323 0.5085
1.500 0.6518 0.01863 0.00916 -0.0815 0.4263 0.6611
1.750 0.6904 0.01820 0.00913 -0.0825 0.4205 1.0000
2.000 0.7158 0.01849 0.00931 -0.0820 0.4145 1.0000
2.250 0.7412 0.01876 0.00947 -0.0814 0.4086 1.0000
2.500 0.7670 0.01903 0.00958 -0.0809 0.4036 1.0000
2.750 0.7926 0.01934 0.00974 -0.0804 0.3987 1.0000
3.000 0.8174 0.01966 0.01003 -0.0798 0.3929 1.0000
3.250 0.8426 0.01997 0.01023 -0.0793 0.3877 1.0000
3.500 0.8684 0.02027 0.01037 -0.0788 0.3835 1.0000
3.750 0.8928 0.02065 0.01072 -0.0782 0.3785 1.0000
4.000 0.9171 0.02101 0.01106 -0.0776 0.3733 1.0000
4.250 0.9419 0.02134 0.01130 -0.0770 0.3687 1.0000
4.500 0.9677 0.02168 0.01149 -0.0766 0.3650 1.0000
4.750 0.9907 0.02213 0.01198 -0.0759 0.3602 1.0000
5.000 1.0142 0.02255 0.01240 -0.0752 0.3556 1.0000
5.250 1.0383 0.02293 0.01271 -0.0746 0.3514 1.0000
5.500 1.0634 0.02329 0.01296 -0.0741 0.3478 1.0000
5.750 1.0855 0.02380 0.01351 -0.0733 0.3435 1.0000
6.000 1.1075 0.02430 0.01404 -0.0725 0.3391 1.0000
6.250 1.1304 0.02476 0.01448 -0.0718 0.3354 1.0000
6.500 1.1543 0.02517 0.01483 -0.0712 0.3320 1.0000
6.750 1.1777 0.02566 0.01525 -0.0706 0.3287 1.0000
7.000 1.1965 0.02629 0.01600 -0.0695 0.3245 1.0000
7.250 1.2168 0.02686 0.01660 -0.0685 0.3205 1.0000
7.500 1.2384 0.02736 0.01709 -0.0677 0.3171 1.0000
7.750 1.2617 0.02783 0.01750 -0.0671 0.3143 1.0000
8.000 1.2823 0.02846 0.01814 -0.0662 0.3113 1.0000
8.250 1.2972 0.02926 0.01908 -0.0647 0.3076 1.0000
8.500 1.3139 0.02996 0.01984 -0.0634 0.3040 1.0000
8.750 1.3328 0.03056 0.02045 -0.0624 0.3008 1.0000
9.000 1.3541 0.03107 0.02093 -0.0616 0.2981 1.0000
9.250 1.3748 0.03171 0.02154 -0.0609 0.2956 1.0000
9.500 1.3815 0.03282 0.02283 -0.0586 0.2924 1.0000
9.750 1.3887 0.03383 0.02396 -0.0562 0.2893 1.0000
10.000 1.3988 0.03471 0.02491 -0.0543 0.2863 1.0000
10.250 1.4135 0.03544 0.02565 -0.0530 0.2836 1.0000
10.500 1.4336 0.03601 0.02619 -0.0522 0.2812 1.0000
10.750 1.4434 0.03708 0.02731 -0.0506 0.2788 1.0000
11.000 1.4336 0.03907 0.02953 -0.0476 0.2759 1.0000
11.250 1.4275 0.04112 0.03174 -0.0454 0.2730 1.0000
11.500 1.4277 0.04294 0.03366 -0.0440 0.2703 1.0000
11.750 1.4370 0.04422 0.03498 -0.0430 0.2678 1.0000
12.000 1.4570 0.04482 0.03556 -0.0425 0.2656 1.0000
12.250 1.4647 0.04636 0.03714 -0.0416 0.2633 1.0000
12.500 1.3747 0.05656 0.04784 -0.0409 0.2584 1.0000
12.750 1.2931 0.06889 0.06045 -0.0436 0.2522 1.0000
13.000 1.3196 0.06853 0.06008 -0.0428 0.2508 1.0000
13.250 1.3565 0.06692 0.05843 -0.0416 0.2498 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 1210 AIRFOIL (e1210-il)