EPPLER 328 AIRFOIL (e328-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 328 AIRFOIL (e328-il) Reynolds number: 200,000 Max Cl/Cd: 63.5 at α=9.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e328-il-200000.txt Download as CSV file: xf-e328-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 328 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3890 0.08964 0.08630 -0.0366 1.0000 0.0373
-9.000 -0.4024 0.08461 0.08127 -0.0411 1.0000 0.0374
-8.750 -0.4235 0.08096 0.07759 -0.0420 1.0000 0.0375
-8.500 -0.4418 0.07849 0.07510 -0.0398 1.0000 0.0376
-8.250 -0.4607 0.07608 0.07261 -0.0370 1.0000 0.0377
-8.000 -0.4757 0.07371 0.07012 -0.0341 1.0000 0.0377
-7.750 -0.4725 0.06832 0.06487 -0.0333 1.0000 0.0383
-7.500 -0.4680 0.06538 0.06204 -0.0316 1.0000 0.0391
-7.250 -0.4722 0.06309 0.05974 -0.0289 1.0000 0.0393
-7.000 -0.4779 0.06110 0.05780 -0.0256 1.0000 0.0401
-6.750 -0.4543 0.05713 0.05370 -0.0290 0.9902 0.0421
-6.500 -0.4294 0.05522 0.05091 -0.0311 0.9645 0.0465
-6.250 -0.4060 0.04782 0.04369 -0.0345 0.9484 0.0481
-6.000 -0.3729 0.04438 0.04019 -0.0377 0.9285 0.0508
-5.750 -0.3010 0.02377 0.01920 -0.0434 0.8438 0.0593
-5.500 -0.2858 0.02165 0.01692 -0.0421 0.8198 0.0619
-5.250 -0.2843 0.03573 0.03049 -0.0419 0.8470 0.0658
-5.000 -0.2683 0.03339 0.02773 -0.0401 0.8210 0.0724
-4.750 -0.2485 0.03217 0.02629 -0.0386 0.7966 0.0794
-4.500 -0.2320 0.03036 0.02429 -0.0371 0.7757 0.0914
-4.250 -0.1970 0.02366 0.01616 -0.0323 0.7620 0.0363
-4.000 -0.1692 0.02096 0.01291 -0.0309 0.7452 0.0318
-3.750 -0.1418 0.01993 0.01155 -0.0299 0.7284 0.0308
-3.500 -0.1138 0.01862 0.01006 -0.0293 0.7124 0.0308
-3.250 -0.0872 0.01734 0.00872 -0.0288 0.6973 0.0317
-3.000 -0.0630 0.01654 0.00788 -0.0278 0.6831 0.0332
-2.750 -0.0399 0.01595 0.00719 -0.0265 0.6699 0.0371
-2.500 -0.0188 0.01533 0.00655 -0.0251 0.6577 0.0444
-2.250 0.0000 0.01459 0.00582 -0.0232 0.6452 0.0605
-2.000 0.0114 0.01326 0.00512 -0.0202 0.6342 0.1938
-1.750 -0.0016 0.01198 0.00626 -0.0101 0.6267 0.8285
-1.500 0.0347 0.01391 0.00805 -0.0073 0.6143 0.8867
-1.250 0.1982 0.01627 0.00980 -0.0294 0.5947 0.9338
-1.000 0.2734 0.01618 0.00947 -0.0381 0.5807 0.9517
-0.750 0.3050 0.01605 0.00920 -0.0392 0.5700 0.9567
-0.500 0.3284 0.01608 0.00906 -0.0386 0.5617 0.9617
-0.250 0.3638 0.01581 0.00869 -0.0404 0.5515 0.9646
0.000 0.3951 0.01571 0.00845 -0.0414 0.5434 0.9681
0.250 0.4213 0.01566 0.00833 -0.0414 0.5349 0.9718
0.500 0.4490 0.01562 0.00817 -0.0418 0.5272 0.9750
0.750 0.4810 0.01545 0.00794 -0.0430 0.5188 0.9774
1.000 0.5105 0.01541 0.00779 -0.0437 0.5120 0.9801
1.250 0.5374 0.01538 0.00773 -0.0438 0.5044 0.9828
1.500 0.5637 0.01544 0.00767 -0.0439 0.4985 0.9853
1.750 0.5944 0.01530 0.00755 -0.0449 0.4910 0.9872
2.000 0.6242 0.01526 0.00743 -0.0456 0.4848 0.9893
2.250 0.6519 0.01527 0.00744 -0.0460 0.4783 0.9912
2.500 0.6793 0.01529 0.00743 -0.0463 0.4721 0.9932
2.750 0.7074 0.01538 0.00742 -0.0467 0.4670 0.9949
3.000 0.7350 0.01534 0.00747 -0.0471 0.4602 0.9963
3.250 0.7639 0.01536 0.00744 -0.0477 0.4546 0.9980
3.500 0.7922 0.01545 0.00751 -0.0482 0.4492 0.9995
3.750 0.8159 0.01553 0.00766 -0.0478 0.4431 1.0000
4.000 0.8383 0.01567 0.00776 -0.0470 0.4382 1.0000
4.250 0.8603 0.01586 0.00797 -0.0462 0.4333 1.0000
4.500 0.8817 0.01600 0.00817 -0.0453 0.4274 1.0000
4.750 0.9041 0.01615 0.00831 -0.0446 0.4225 1.0000
5.000 0.9255 0.01636 0.00855 -0.0437 0.4175 1.0000
5.250 0.9464 0.01650 0.00878 -0.0427 0.4116 1.0000
5.500 0.9686 0.01667 0.00893 -0.0418 0.4068 1.0000
5.750 0.9893 0.01689 0.00922 -0.0408 0.4015 1.0000
6.000 1.0096 0.01705 0.00946 -0.0397 0.3956 1.0000
6.250 1.0315 0.01723 0.00961 -0.0389 0.3907 1.0000
6.500 1.0508 0.01746 0.00996 -0.0376 0.3851 1.0000
6.750 1.0705 0.01759 0.01016 -0.0364 0.3789 1.0000
7.000 1.0924 0.01781 0.01032 -0.0355 0.3739 1.0000
7.250 1.1090 0.01800 0.01071 -0.0338 0.3673 1.0000
7.500 1.1287 0.01815 0.01087 -0.0325 0.3615 1.0000
7.750 1.1467 0.01839 0.01119 -0.0310 0.3557 1.0000
8.000 1.1634 0.01856 0.01147 -0.0293 0.3490 1.0000
8.250 1.1829 0.01877 0.01163 -0.0280 0.3433 1.0000
8.500 1.1958 0.01896 0.01203 -0.0256 0.3360 1.0000
8.750 1.2127 0.01911 0.01216 -0.0239 0.3297 1.0000
9.000 1.2248 0.01936 0.01258 -0.0214 0.3226 1.0000
9.250 1.2383 0.01950 0.01276 -0.0190 0.3159 1.0000
9.500 1.2485 0.01974 0.01311 -0.0162 0.3087 1.0000
9.750 1.2579 0.01989 0.01332 -0.0132 0.3015 1.0000
10.000 1.2642 0.02011 0.01363 -0.0096 0.2942 1.0000
10.250 1.2684 0.02026 0.01386 -0.0056 0.2871 1.0000
10.500 1.2699 0.02047 0.01414 -0.0012 0.2800 1.0000
10.750 1.2721 0.02064 0.01435 0.0030 0.2722 1.0000
11.000 1.2785 0.02094 0.01475 0.0062 0.2630 1.0000
11.250 1.2879 0.02132 0.01514 0.0087 0.2535 1.0000
11.500 1.2947 0.02176 0.01562 0.0116 0.2430 1.0000
11.750 1.2997 0.02232 0.01629 0.0146 0.2321 1.0000
12.000 1.3040 0.02304 0.01706 0.0173 0.2209 1.0000
12.250 1.3070 0.02395 0.01800 0.0199 0.2096 1.0000
12.500 1.3075 0.02512 0.01917 0.0223 0.1983 1.0000
12.750 1.3066 0.02656 0.02063 0.0243 0.1869 1.0000
13.000 1.3051 0.02825 0.02238 0.0259 0.1755 1.0000
13.250 1.3019 0.03029 0.02446 0.0270 0.1646 1.0000
13.500 1.2971 0.03265 0.02685 0.0279 0.1550 1.0000
13.750 1.2890 0.03548 0.02964 0.0284 0.1462 1.0000
14.000 1.2841 0.03828 0.03255 0.0285 0.1368 1.0000
14.250 1.2761 0.04147 0.03577 0.0285 0.1286 1.0000
14.500 1.2668 0.04490 0.03919 0.0282 0.1216 1.0000
14.750 1.2593 0.04833 0.04270 0.0277 0.1140 1.0000
15.000 1.2488 0.05204 0.04636 0.0271 0.1082 1.0000
15.250 1.2422 0.05561 0.05007 0.0263 0.1017 1.0000
15.500 1.2325 0.05950 0.05389 0.0255 0.0962 1.0000
15.750 1.2263 0.06330 0.05783 0.0244 0.0905 1.0000
16.000 1.2179 0.06731 0.06181 0.0232 0.0854 1.0000
16.250 1.2122 0.07117 0.06577 0.0220 0.0804 1.0000
16.500 1.2051 0.07530 0.06994 0.0205 0.0757 1.0000
16.750 1.1998 0.07914 0.07378 0.0194 0.0712 1.0000
17.000 1.1938 0.08338 0.07813 0.0177 0.0670 1.0000
17.250 1.1922 0.08653 0.08116 0.0169 0.0627 1.0000
17.500 1.1855 0.09115 0.08600 0.0149 0.0596 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 328 AIRFOIL (e328-il)