GOE 513 AIRFOIL (goe513-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 513 AIRFOIL (goe513-il) Reynolds number: 100,000 Max Cl/Cd: 43.04 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe513-il-100000-n5.txt Download as CSV file: xf-goe513-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 513 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
------ -------- --------- --------- -------- -------- --------
-12.000 -0.3114 0.13317 0.12815 -0.0370 0.9996 0.0597
-11.750 -0.2927 0.12928 0.12425 -0.0397 0.9966 0.0578
-11.250 -0.3040 0.11379 0.10871 -0.0514 0.9896 0.0532
-11.000 -0.2968 0.10900 0.10393 -0.0550 0.9859 0.0530
-10.750 -0.2873 0.10447 0.09938 -0.0587 0.9826 0.0528
-10.500 -0.2814 0.09989 0.09481 -0.0619 0.9781 0.0524
-10.250 -0.2805 0.09451 0.08943 -0.0658 0.9733 0.0520
-10.000 -0.2830 0.08776 0.08268 -0.0715 0.9693 0.0515
-9.750 -0.3067 0.07896 0.07390 -0.0773 0.9610 0.0508
-9.500 -0.4421 0.05780 0.05210 -0.0860 0.9420 0.0487
-9.250 -0.4655 0.05380 0.04777 -0.0825 0.9298 0.0488
-9.000 -0.4729 0.05012 0.04372 -0.0802 0.9217 0.0490
-8.750 -0.4769 0.04777 0.04111 -0.0767 0.9120 0.0495
-8.500 -0.4763 0.04537 0.03839 -0.0737 0.9040 0.0500
-8.250 -0.4716 0.04310 0.03577 -0.0709 0.8956 0.0506
-8.000 -0.4657 0.04112 0.03347 -0.0678 0.8873 0.0510
-7.750 -0.4545 0.03921 0.03122 -0.0655 0.8800 0.0513
-7.500 -0.4373 0.03739 0.02908 -0.0639 0.8738 0.0517
-7.250 -0.4257 0.03596 0.02737 -0.0611 0.8648 0.0521
-7.000 -0.3996 0.03434 0.02543 -0.0608 0.8609 0.0526
-6.750 -0.3896 0.03338 0.02425 -0.0575 0.8512 0.0530
-6.500 -0.3620 0.03203 0.02279 -0.0574 0.8465 0.0537
-6.250 -0.3371 0.03096 0.02171 -0.0569 0.8409 0.0549
-6.000 -0.3184 0.03012 0.02080 -0.0551 0.8325 0.0558
-5.750 -0.2870 0.02901 0.01958 -0.0555 0.8284 0.0573
-5.500 -0.2688 0.02828 0.01875 -0.0534 0.8194 0.0583
-5.250 -0.2398 0.02732 0.01768 -0.0533 0.8135 0.0592
-5.000 -0.2040 0.02629 0.01655 -0.0543 0.8097 0.0604
-4.750 -0.1902 0.02572 0.01600 -0.0515 0.7984 0.0613
-4.500 -0.1562 0.02477 0.01503 -0.0522 0.7932 0.0630
-4.250 -0.1376 0.02427 0.01449 -0.0502 0.7824 0.0648
-4.000 -0.1022 0.02348 0.01359 -0.0511 0.7759 0.0678
-3.750 -0.0793 0.02291 0.01300 -0.0499 0.7653 0.0706
-3.500 -0.0420 0.02214 0.01216 -0.0513 0.7574 0.0749
-3.250 -0.0149 0.02160 0.01159 -0.0508 0.7458 0.0797
-3.000 0.0311 0.02071 0.01071 -0.0541 0.7372 0.0932
-2.750 0.0588 0.02002 0.01043 -0.0540 0.7237 0.1407
-2.500 0.0919 0.01991 0.01039 -0.0547 0.7112 0.2119
-2.250 0.1263 0.01990 0.01022 -0.0556 0.6981 0.2370
-2.000 0.1432 0.02003 0.01027 -0.0531 0.6822 0.2553
-1.750 0.1620 0.02015 0.01026 -0.0510 0.6666 0.2715
-1.500 0.1829 0.02017 0.01023 -0.0493 0.6512 0.2846
-1.250 0.2046 0.02017 0.01015 -0.0479 0.6361 0.2988
-1.000 0.2291 0.02014 0.00999 -0.0470 0.6218 0.3114
-0.750 0.2552 0.02005 0.00981 -0.0465 0.6076 0.3199
-0.500 0.2831 0.02000 0.00958 -0.0464 0.5943 0.3255
-0.250 0.3171 0.01987 0.00938 -0.0475 0.5820 0.3302
0.000 0.3531 0.01981 0.00917 -0.0491 0.5702 0.3355
0.250 0.3830 0.01984 0.00908 -0.0494 0.5581 0.3411
0.500 0.4152 0.01987 0.00900 -0.0503 0.5463 0.3471
0.750 0.4494 0.01993 0.00893 -0.0516 0.5355 0.3548
1.000 0.4746 0.02004 0.00898 -0.0511 0.5252 0.3625
1.250 0.5063 0.02010 0.00894 -0.0519 0.5164 0.3708
1.500 0.5277 0.02020 0.00903 -0.0507 0.5079 0.3798
1.750 0.5522 0.02025 0.00907 -0.0502 0.5000 0.3912
2.000 0.5758 0.02031 0.00912 -0.0495 0.4928 0.4075
2.500 0.7741 0.02082 0.01097 -0.0784 0.4674 0.9434
2.750 0.8097 0.02129 0.01131 -0.0800 0.4598 0.9626
3.000 0.8500 0.02172 0.01168 -0.0827 0.4510 0.9776
3.250 0.8964 0.02204 0.01191 -0.0868 0.4416 0.9888
3.500 0.9447 0.02228 0.01210 -0.0915 0.4318 0.9984
3.750 0.9653 0.02251 0.01227 -0.0903 0.4240 1.0000
4.000 0.9783 0.02275 0.01245 -0.0876 0.4174 1.0000
4.250 0.9888 0.02300 0.01272 -0.0845 0.4099 1.0000
4.500 1.0007 0.02325 0.01289 -0.0816 0.4036 1.0000
4.750 1.0106 0.02352 0.01316 -0.0784 0.3967 1.0000
5.000 1.0183 0.02377 0.01339 -0.0747 0.3897 1.0000
5.250 1.0277 0.02403 0.01356 -0.0713 0.3838 1.0000
5.500 1.0345 0.02433 0.01390 -0.0675 0.3763 1.0000
5.750 1.0425 0.02463 0.01414 -0.0640 0.3698 1.0000
6.000 1.0515 0.02498 0.01446 -0.0607 0.3633 1.0000
6.250 1.0598 0.02535 0.01484 -0.0574 0.3564 1.0000
6.500 1.0695 0.02571 0.01512 -0.0543 0.3507 1.0000
6.750 1.0791 0.02616 0.01559 -0.0514 0.3443 1.0000
7.000 1.0888 0.02660 0.01602 -0.0484 0.3385 1.0000
7.250 1.1004 0.02704 0.01639 -0.0459 0.3337 1.0000
7.500 1.1105 0.02757 0.01696 -0.0432 0.3281 1.0000
7.750 1.1207 0.02810 0.01750 -0.0405 0.3227 1.0000
8.000 1.1322 0.02861 0.01794 -0.0381 0.3181 1.0000
8.250 1.1432 0.02920 0.01853 -0.0357 0.3135 1.0000
8.500 1.1527 0.02983 0.01922 -0.0331 0.3083 1.0000
8.750 1.1635 0.03044 0.01982 -0.0308 0.3040 1.0000
9.000 1.1776 0.03100 0.02030 -0.0290 0.3004 1.0000
9.250 1.1883 0.03172 0.02110 -0.0268 0.2966 1.0000
9.500 1.1985 0.03248 0.02192 -0.0247 0.2928 1.0000
9.750 1.2097 0.03321 0.02268 -0.0227 0.2893 1.0000
10.000 1.2224 0.03389 0.02334 -0.0209 0.2860 1.0000
10.250 1.2392 0.03447 0.02385 -0.0197 0.2830 1.0000
10.500 1.2457 0.03547 0.02500 -0.0173 0.2798 1.0000
10.750 1.2527 0.03647 0.02609 -0.0151 0.2763 1.0000
11.000 1.2622 0.03739 0.02707 -0.0132 0.2732 1.0000
11.250 1.2741 0.03824 0.02795 -0.0117 0.2707 1.0000
11.500 1.2890 0.03899 0.02870 -0.0105 0.2683 1.0000
11.750 1.3069 0.03967 0.02935 -0.0096 0.2660 1.0000
12.000 1.3070 0.04116 0.03101 -0.0072 0.2634 1.0000
12.250 1.3084 0.04265 0.03265 -0.0049 0.2606 1.0000
12.500 1.3120 0.04406 0.03418 -0.0030 0.2580 1.0000
12.750 1.3185 0.04534 0.03553 -0.0015 0.2554 1.0000
13.000 1.3292 0.04641 0.03664 -0.0003 0.2532 1.0000
13.250 1.3442 0.04728 0.03753 0.0006 0.2513 1.0000
13.500 1.3614 0.04809 0.03835 0.0013 0.2493 1.0000
13.750 1.3442 0.05103 0.04154 0.0040 0.2470 1.0000
14.000 1.3275 0.05423 0.04496 0.0062 0.2445 1.0000
14.250 1.3133 0.05747 0.04837 0.0078 0.2419 1.0000
14.500 1.3053 0.06034 0.05137 0.0090 0.2393 1.0000
14.750 1.3043 0.06270 0.05382 0.0099 0.2373 1.0000
15.000 1.3146 0.06404 0.05521 0.0105 0.2355 1.0000
15.250 1.3370 0.06427 0.05543 0.0110 0.2338 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 513 AIRFOIL (goe513-il)