EPPLER 340 AIRFOIL (e340-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 340 AIRFOIL (e340-il) Reynolds number: 100,000 Max Cl/Cd: 42.07 at α=9.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e340-il-100000-n5.txt Download as CSV file: xf-e340-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 340 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5433 0.08773 0.08324 -0.0113 1.0000 0.0255
-9.500 -0.5443 0.08434 0.07988 -0.0120 1.0000 0.0251
-9.250 -0.5514 0.08019 0.07572 -0.0132 1.0000 0.0247
-9.000 -0.5616 0.07605 0.07153 -0.0137 1.0000 0.0243
-8.750 -0.5733 0.07208 0.06750 -0.0132 1.0000 0.0238
-8.250 -0.6104 0.06108 0.05595 -0.0081 1.0000 0.0211
-8.000 -0.6101 0.05806 0.05282 -0.0060 1.0000 0.0209
-7.750 -0.6084 0.05502 0.04964 -0.0039 1.0000 0.0207
-7.500 -0.6056 0.05189 0.04633 -0.0015 1.0000 0.0204
-7.250 -0.6015 0.04872 0.04295 0.0011 1.0000 0.0202
-7.000 -0.5832 0.04472 0.03850 0.0012 0.8991 0.0199
-6.750 -0.5682 0.04141 0.03466 0.0030 0.8490 0.0196
-6.500 -0.5558 0.03856 0.03127 0.0059 0.8136 0.0194
-6.250 -0.5412 0.03594 0.02812 0.0085 0.7855 0.0193
-6.000 -0.5234 0.03356 0.02523 0.0108 0.7617 0.0194
-5.750 -0.5023 0.03141 0.02260 0.0125 0.7408 0.0195
-5.500 -0.4782 0.02941 0.02016 0.0138 0.7212 0.0199
-5.250 -0.4515 0.02768 0.01794 0.0146 0.7030 0.0210
-5.000 -0.4263 0.02656 0.01669 0.0152 0.6859 0.0225
-4.750 -0.3989 0.02543 0.01533 0.0156 0.6699 0.0242
-4.500 -0.3696 0.02412 0.01376 0.0159 0.6550 0.0256
-4.250 -0.3435 0.02306 0.01263 0.0163 0.6411 0.0271
-4.000 -0.3183 0.02214 0.01160 0.0171 0.6274 0.0292
-3.750 -0.2947 0.02140 0.01080 0.0180 0.6146 0.0330
-3.250 -0.2490 0.02004 0.00923 0.0202 0.5922 0.0450
-3.000 -0.2256 0.01938 0.00858 0.0212 0.5814 0.0622
-2.750 -0.2022 0.01850 0.00797 0.0221 0.5720 0.1191
-2.500 -0.1293 0.01866 0.01112 0.0186 0.5582 0.8559
-2.250 -0.0948 0.02004 0.01221 0.0195 0.5476 0.8964
-2.000 -0.0399 0.02112 0.01293 0.0161 0.5371 0.9278
-1.750 0.0151 0.02128 0.01280 0.0114 0.5260 0.9442
-1.500 0.0491 0.02115 0.01243 0.0101 0.5172 0.9497
-1.250 0.0776 0.02105 0.01214 0.0097 0.5091 0.9548
-1.000 0.1113 0.02088 0.01177 0.0082 0.5013 0.9582
-0.750 0.1399 0.02081 0.01154 0.0078 0.4940 0.9629
-0.500 0.1720 0.02067 0.01121 0.0066 0.4874 0.9660
-0.250 0.2039 0.02054 0.01097 0.0055 0.4796 0.9694
0.000 0.2322 0.02050 0.01077 0.0051 0.4737 0.9731
0.250 0.2653 0.02037 0.01055 0.0037 0.4671 0.9756
0.500 0.2960 0.02030 0.01038 0.0027 0.4610 0.9785
0.750 0.3251 0.02028 0.01021 0.0022 0.4560 0.9814
1.000 0.3570 0.02020 0.01011 0.0009 0.4494 0.9837
1.250 0.3872 0.02016 0.01000 0.0001 0.4438 0.9863
1.500 0.4165 0.02017 0.00988 -0.0006 0.4395 0.9887
1.750 0.4480 0.02013 0.00986 -0.0018 0.4335 0.9906
2.000 0.4783 0.02012 0.00981 -0.0026 0.4282 0.9928
2.250 0.5076 0.02014 0.00975 -0.0033 0.4239 0.9949
2.500 0.5386 0.02015 0.00979 -0.0044 0.4190 0.9965
2.750 0.5693 0.02019 0.00984 -0.0054 0.4139 0.9983
3.000 0.5988 0.02023 0.00985 -0.0061 0.4095 1.0000
3.250 0.6232 0.02038 0.00994 -0.0058 0.4058 1.0000
3.500 0.6478 0.02057 0.01025 -0.0057 0.4006 1.0000
3.750 0.6722 0.02075 0.01045 -0.0054 0.3962 1.0000
4.000 0.6965 0.02091 0.01058 -0.0050 0.3923 1.0000
4.250 0.7207 0.02113 0.01081 -0.0047 0.3885 1.0000
4.500 0.7448 0.02140 0.01122 -0.0045 0.3838 1.0000
4.750 0.7688 0.02163 0.01150 -0.0041 0.3796 1.0000
5.000 0.7927 0.02183 0.01169 -0.0037 0.3759 1.0000
5.250 0.8164 0.02210 0.01204 -0.0033 0.3719 1.0000
5.500 0.8397 0.02244 0.01252 -0.0030 0.3670 1.0000
5.750 0.8630 0.02270 0.01284 -0.0025 0.3627 1.0000
6.000 0.8866 0.02291 0.01307 -0.0020 0.3592 1.0000
6.250 0.9093 0.02330 0.01359 -0.0015 0.3548 1.0000
6.500 0.9315 0.02369 0.01414 -0.0010 0.3498 1.0000
6.750 0.9541 0.02395 0.01447 -0.0004 0.3456 1.0000
7.000 0.9773 0.02416 0.01467 0.0002 0.3422 1.0000
7.250 0.9975 0.02471 0.01548 0.0008 0.3364 1.0000
7.500 1.0188 0.02505 0.01596 0.0016 0.3314 1.0000
7.750 1.0411 0.02526 0.01618 0.0023 0.3274 1.0000
8.000 1.0603 0.02579 0.01692 0.0032 0.3220 1.0000
8.250 1.0797 0.02622 0.01752 0.0042 0.3164 1.0000
8.500 1.1011 0.02639 0.01772 0.0051 0.3119 1.0000
8.750 1.1181 0.02698 0.01853 0.0063 0.3059 1.0000
9.000 1.1359 0.02738 0.01908 0.0075 0.2999 1.0000
9.250 1.1566 0.02749 0.01921 0.0085 0.2953 1.0000
9.500 1.1690 0.02829 0.02031 0.0102 0.2881 1.0000
9.750 1.1862 0.02855 0.02066 0.0116 0.2822 1.0000
10.000 1.1990 0.02914 0.02144 0.0134 0.2755 1.0000
10.250 1.2119 0.02961 0.02206 0.0153 0.2685 1.0000
10.500 1.2233 0.03012 0.02271 0.0174 0.2619 1.0000
10.750 1.2312 0.03079 0.02354 0.0198 0.2545 1.0000
11.000 1.2386 0.03137 0.02423 0.0223 0.2477 1.0000
11.250 1.2400 0.03223 0.02525 0.0254 0.2402 1.0000
11.500 1.2405 0.03298 0.02607 0.0287 0.2335 1.0000
11.750 1.2306 0.03413 0.02738 0.0331 0.2270 1.0000
12.000 1.2174 0.03495 0.02820 0.0381 0.2223 1.0000
12.250 1.1924 0.03678 0.03016 0.0430 0.2177 1.0000
12.500 1.1695 0.03903 0.03250 0.0466 0.2127 1.0000
12.750 1.1579 0.04075 0.03420 0.0491 0.2073 1.0000
13.000 1.1287 0.04448 0.03807 0.0509 0.2022 1.0000
13.250 1.1099 0.04758 0.04121 0.0522 0.1964 1.0000
13.500 1.0927 0.05084 0.04450 0.0531 0.1906 1.0000
13.750 1.0669 0.05548 0.04922 0.0531 0.1846 1.0000
14.000 1.0601 0.05805 0.05176 0.0535 0.1779 1.0000
14.250 1.0347 0.06338 0.05719 0.0525 0.1715 1.0000
14.500 1.0344 0.06544 0.05921 0.0526 0.1646 1.0000
14.750 1.0112 0.07104 0.06492 0.0511 0.1580 1.0000
15.000 1.0133 0.07307 0.06691 0.0510 0.1510 1.0000
15.250 0.9940 0.07861 0.07255 0.0492 0.1445 1.0000
15.500 0.9980 0.08056 0.07445 0.0489 0.1378 1.0000
15.750 0.9810 0.08609 0.08011 0.0470 0.1315 1.0000
16.000 0.9864 0.08793 0.08187 0.0467 0.1250 1.0000
16.250 0.9697 0.09369 0.08777 0.0445 0.1193 1.0000
16.500 0.9769 0.09532 0.08934 0.0441 0.1130 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 340 AIRFOIL (e340-il)