EPPLER 598 AIRFOIL (e598-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 598 AIRFOIL (e598-il) Reynolds number: 200,000 Max Cl/Cd: 72.17 at α=9.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e598-il-200000.txt Download as CSV file: xf-e598-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 598 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 -0.3249 0.10706 0.10364 -0.0081 0.8022 0.0404
-9.000 -0.3299 0.10418 0.10069 -0.0116 0.7849 0.0414
-8.750 -0.3372 0.10089 0.09737 -0.0160 0.7710 0.0417
-8.500 -0.3248 0.09694 0.09335 -0.0150 0.7578 0.0422
-8.250 -0.3116 0.09410 0.09042 -0.0142 0.7459 0.0430
-8.000 -0.3036 0.09132 0.08759 -0.0148 0.7349 0.0440
-7.750 -0.2986 0.08835 0.08459 -0.0161 0.7253 0.0449
-7.500 -0.2963 0.08529 0.08150 -0.0178 0.7171 0.0463
-7.250 -0.3040 0.08197 0.07821 -0.0216 0.7090 0.0479
-7.000 -0.3121 0.07642 0.07260 -0.0333 0.7031 0.0486
-6.750 -0.3021 0.07296 0.06913 -0.0314 0.6962 0.0492
-6.500 -0.2895 0.07032 0.06647 -0.0307 0.6886 0.0499
-6.250 -0.2777 0.06749 0.06356 -0.0316 0.6827 0.0511
-6.000 -0.2649 0.06391 0.05994 -0.0344 0.6759 0.0526
-5.750 -0.2782 0.03501 0.02968 -0.0493 0.6786 0.0263
-5.500 -0.2595 0.03120 0.02550 -0.0494 0.6721 0.0257
-5.250 -0.2405 0.02683 0.02040 -0.0490 0.6669 0.0254
-5.000 -0.2176 0.02356 0.01642 -0.0484 0.6609 0.0257
-4.750 -0.1927 0.02162 0.01410 -0.0479 0.6548 0.0271
-4.500 -0.1666 0.02077 0.01298 -0.0475 0.6496 0.0298
-4.250 -0.1398 0.01989 0.01201 -0.0473 0.6433 0.0336
-4.000 -0.1129 0.01869 0.01059 -0.0467 0.6375 0.0378
-3.750 -0.0860 0.01809 0.00984 -0.0463 0.6328 0.0454
-3.500 -0.0580 0.01768 0.00932 -0.0460 0.6270 0.0568
-3.250 -0.0309 0.01727 0.00894 -0.0459 0.6214 0.0686
-3.000 -0.0036 0.01702 0.00861 -0.0455 0.6168 0.0804
-2.750 0.0241 0.01668 0.00822 -0.0454 0.6116 0.0919
-2.500 0.0520 0.01637 0.00784 -0.0452 0.6063 0.1033
-2.250 0.0795 0.01613 0.00753 -0.0449 0.6017 0.1150
-2.000 0.1067 0.01579 0.00718 -0.0446 0.5973 0.1275
-1.750 0.1344 0.01545 0.00689 -0.0445 0.5919 0.1411
-1.500 0.1618 0.01516 0.00659 -0.0442 0.5873 0.1562
-1.250 0.1889 0.01493 0.00632 -0.0439 0.5833 0.1737
-1.000 0.2163 0.01473 0.00619 -0.0437 0.5783 0.1952
-0.750 0.2434 0.01446 0.00605 -0.0436 0.5733 0.2231
-0.500 0.2700 0.01416 0.00589 -0.0432 0.5692 0.2689
-0.250 0.2934 0.01347 0.00578 -0.0426 0.5656 0.4271
0.000 0.3655 0.01205 0.00574 -0.0507 0.5588 1.0000
0.250 0.3918 0.01218 0.00570 -0.0503 0.5544 1.0000
0.500 0.4180 0.01232 0.00566 -0.0497 0.5506 1.0000
0.750 0.4443 0.01251 0.00577 -0.0494 0.5459 1.0000
1.000 0.4707 0.01266 0.00586 -0.0491 0.5408 1.0000
1.250 0.4970 0.01280 0.00587 -0.0486 0.5366 1.0000
1.500 0.5234 0.01299 0.00590 -0.0481 0.5331 1.0000
1.750 0.5497 0.01318 0.00611 -0.0479 0.5277 1.0000
2.000 0.5762 0.01334 0.00622 -0.0475 0.5229 1.0000
2.250 0.6027 0.01348 0.00625 -0.0471 0.5189 1.0000
2.500 0.6291 0.01370 0.00641 -0.0468 0.5147 1.0000
2.750 0.6554 0.01390 0.00664 -0.0465 0.5093 1.0000
3.000 0.6820 0.01404 0.00673 -0.0462 0.5048 1.0000
3.250 0.7088 0.01421 0.00678 -0.0458 0.5011 1.0000
3.500 0.7349 0.01446 0.00709 -0.0456 0.4957 1.0000
3.750 0.7613 0.01463 0.00726 -0.0453 0.4905 1.0000
4.000 0.7882 0.01476 0.00732 -0.0450 0.4865 1.0000
4.250 0.8145 0.01501 0.00757 -0.0447 0.4819 1.0000
4.500 0.8406 0.01521 0.00783 -0.0445 0.4761 1.0000
4.750 0.8674 0.01534 0.00791 -0.0442 0.4716 1.0000
5.000 0.8942 0.01554 0.00807 -0.0439 0.4675 1.0000
5.250 0.9197 0.01578 0.00844 -0.0437 0.4611 1.0000
5.500 0.9464 0.01590 0.00854 -0.0434 0.4562 1.0000
5.750 0.9734 0.01607 0.00864 -0.0431 0.4521 1.0000
6.000 0.9984 0.01633 0.00906 -0.0429 0.4453 1.0000
6.250 1.0251 0.01643 0.00916 -0.0426 0.4402 1.0000
6.500 1.0515 0.01662 0.00934 -0.0423 0.4352 1.0000
6.750 1.0766 0.01684 0.00969 -0.0420 0.4284 1.0000
7.000 1.1035 0.01690 0.00971 -0.0417 0.4232 1.0000
7.250 1.1284 0.01715 0.01008 -0.0414 0.4164 1.0000
7.500 1.1542 0.01723 0.01020 -0.0411 0.4098 1.0000
7.750 1.1799 0.01740 0.01040 -0.0407 0.4035 1.0000
8.000 1.2046 0.01755 0.01064 -0.0403 0.3957 1.0000
8.250 1.2303 0.01769 0.01078 -0.0400 0.3891 1.0000
8.500 1.2543 0.01788 0.01110 -0.0395 0.3806 1.0000
8.750 1.2791 0.01806 0.01130 -0.0391 0.3732 1.0000
9.000 1.3026 0.01825 0.01160 -0.0386 0.3643 1.0000
9.250 1.3259 0.01851 0.01194 -0.0381 0.3554 1.0000
9.500 1.3495 0.01870 0.01213 -0.0375 0.3464 1.0000
9.750 1.3707 0.01903 0.01262 -0.0368 0.3357 1.0000
10.000 1.3922 0.01936 0.01302 -0.0361 0.3254 1.0000
10.250 1.4129 0.01972 0.01339 -0.0353 0.3147 1.0000
10.500 1.4317 0.02016 0.01393 -0.0344 0.3025 1.0000
10.750 1.4489 0.02069 0.01456 -0.0333 0.2895 1.0000
11.000 1.4644 0.02131 0.01525 -0.0321 0.2762 1.0000
11.250 1.4773 0.02204 0.01604 -0.0306 0.2623 1.0000
11.500 1.4867 0.02292 0.01696 -0.0289 0.2482 1.0000
11.750 1.4906 0.02396 0.01804 -0.0265 0.2347 1.0000
12.000 1.4885 0.02530 0.01940 -0.0239 0.2224 1.0000
12.250 1.4848 0.02710 0.02121 -0.0221 0.2103 1.0000
12.500 1.4796 0.02934 0.02346 -0.0209 0.1982 1.0000
12.750 1.4738 0.03192 0.02605 -0.0202 0.1859 1.0000
13.000 1.4678 0.03475 0.02891 -0.0198 0.1739 1.0000
13.250 1.4606 0.03782 0.03201 -0.0196 0.1625 1.0000
13.500 1.4523 0.04112 0.03531 -0.0195 0.1522 1.0000
13.750 1.4419 0.04468 0.03883 -0.0196 0.1431 1.0000
14.000 1.4334 0.04821 0.04239 -0.0198 0.1336 1.0000
14.250 1.4251 0.05177 0.04599 -0.0201 0.1250 1.0000
14.500 1.4138 0.05568 0.04984 -0.0205 0.1177 1.0000
14.750 1.4068 0.05944 0.05368 -0.0212 0.1095 1.0000
15.000 1.3988 0.06337 0.05762 -0.0219 0.1025 1.0000
15.250 1.3910 0.06743 0.06169 -0.0228 0.0960 1.0000
15.500 1.3850 0.07134 0.06567 -0.0237 0.0895 1.0000
15.750 1.3772 0.07555 0.06985 -0.0248 0.0839 1.0000
16.000 1.3722 0.07955 0.07394 -0.0258 0.0782 1.0000
16.250 1.3661 0.08371 0.07808 -0.0270 0.0733 1.0000
16.500 1.3618 0.08774 0.08219 -0.0281 0.0686 1.0000
16.750 1.3567 0.09194 0.08642 -0.0295 0.0644 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 598 AIRFOIL (e598-il)