GOE 575 AIRFOIL (goe575-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 575 AIRFOIL (goe575-il) Reynolds number: 200,000 Max Cl/Cd: 55.09 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe575-il-200000-n5.txt Download as CSV file: xf-goe575-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 575 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.8019 0.05521 0.05083 -0.0472 1.0000 0.0392
-10.250 -0.8476 0.05085 0.04605 -0.0400 0.9965 0.0394
-10.000 -0.8804 0.04588 0.04031 -0.0338 0.9902 0.0400
-9.750 -0.8676 0.04453 0.03892 -0.0323 0.9876 0.0405
-9.500 -0.8519 0.04319 0.03748 -0.0312 0.9840 0.0409
-9.250 -0.8367 0.04182 0.03598 -0.0298 0.9792 0.0415
-9.000 -0.8206 0.04038 0.03435 -0.0286 0.9753 0.0422
-8.750 -0.8104 0.03875 0.03249 -0.0260 0.9703 0.0431
-8.500 -0.7961 0.03701 0.03041 -0.0242 0.9674 0.0437
-8.250 -0.7905 0.03546 0.02856 -0.0202 0.9623 0.0443
-8.000 -0.7752 0.03401 0.02675 -0.0181 0.9590 0.0449
-7.750 -0.7641 0.03284 0.02518 -0.0148 0.9545 0.0456
-7.500 -0.7437 0.03169 0.02400 -0.0137 0.9509 0.0463
-7.250 -0.7181 0.03070 0.02292 -0.0135 0.9482 0.0470
-7.000 -0.6988 0.02976 0.02186 -0.0119 0.9441 0.0475
-6.750 -0.6772 0.02887 0.02085 -0.0106 0.9399 0.0483
-6.500 -0.6487 0.02800 0.01984 -0.0108 0.9372 0.0489
-6.250 -0.6253 0.02714 0.01883 -0.0098 0.9331 0.0496
-6.000 -0.6018 0.02638 0.01793 -0.0088 0.9282 0.0503
-5.750 -0.5718 0.02577 0.01713 -0.0091 0.9252 0.0515
-5.500 -0.5369 0.02489 0.01616 -0.0105 0.9232 0.0524
-5.250 -0.5185 0.02410 0.01539 -0.0085 0.9158 0.0531
-5.000 -0.4862 0.02339 0.01467 -0.0094 0.9124 0.0538
-4.750 -0.4506 0.02269 0.01396 -0.0109 0.9101 0.0549
-4.500 -0.4309 0.02215 0.01339 -0.0091 0.9022 0.0558
-4.250 -0.3984 0.02154 0.01275 -0.0099 0.8979 0.0569
-4.000 -0.3621 0.02097 0.01212 -0.0114 0.8950 0.0583
-3.750 -0.3413 0.02053 0.01164 -0.0097 0.8859 0.0589
-3.500 -0.3088 0.01981 0.01098 -0.0105 0.8810 0.0604
-3.250 -0.2708 0.01914 0.01032 -0.0123 0.8774 0.0619
-3.000 -0.2521 0.01877 0.00993 -0.0102 0.8650 0.0629
-2.750 -0.2189 0.01819 0.00934 -0.0109 0.8576 0.0650
-2.500 -0.1950 0.01778 0.00892 -0.0098 0.8446 0.0672
-2.250 -0.1683 0.01738 0.00855 -0.0093 0.8324 0.0700
-2.000 -0.1320 0.01686 0.00805 -0.0108 0.8244 0.0737
-1.750 -0.1006 0.01644 0.00771 -0.0112 0.8107 0.0810
-1.500 -0.0612 0.01593 0.00731 -0.0134 0.7985 0.1029
-1.250 -0.0144 0.01537 0.00681 -0.0171 0.7860 0.1262
-1.000 0.0392 0.01481 0.00627 -0.0222 0.7715 0.1504
-0.750 0.0989 0.01429 0.00571 -0.0288 0.7531 0.1662
-0.500 0.1570 0.01392 0.00527 -0.0352 0.7314 0.1841
-0.250 0.2061 0.01371 0.00497 -0.0397 0.7085 0.1990
0.000 0.2454 0.01363 0.00481 -0.0421 0.6874 0.2129
0.250 0.2782 0.01357 0.00473 -0.0431 0.6678 0.2335
0.500 0.3071 0.01349 0.00470 -0.0434 0.6491 0.2638
0.750 0.3343 0.01323 0.00471 -0.0434 0.6305 0.3519
1.250 0.6588 0.01471 0.00739 -0.1021 0.5450 0.9955
1.500 0.6966 0.01473 0.00723 -0.1044 0.5200 1.0000
1.750 0.7129 0.01494 0.00727 -0.1019 0.4982 1.0000
2.000 0.7304 0.01513 0.00734 -0.0997 0.4797 1.0000
2.250 0.7477 0.01533 0.00742 -0.0975 0.4614 1.0000
2.500 0.7656 0.01553 0.00751 -0.0954 0.4448 1.0000
2.750 0.7848 0.01571 0.00762 -0.0936 0.4314 1.0000
3.000 0.8030 0.01592 0.00773 -0.0916 0.4170 1.0000
3.250 0.8217 0.01612 0.00786 -0.0897 0.4057 1.0000
3.500 0.8409 0.01632 0.00800 -0.0879 0.3947 1.0000
3.750 0.8603 0.01652 0.00815 -0.0862 0.3849 1.0000
4.000 0.8794 0.01674 0.00832 -0.0843 0.3756 1.0000
4.250 0.8988 0.01695 0.00849 -0.0826 0.3667 1.0000
4.500 0.9177 0.01718 0.00868 -0.0808 0.3579 1.0000
4.750 0.9366 0.01740 0.00888 -0.0790 0.3486 1.0000
5.000 0.9552 0.01765 0.00909 -0.0772 0.3402 1.0000
5.250 0.9740 0.01789 0.00932 -0.0754 0.3319 1.0000
5.500 0.9912 0.01816 0.00955 -0.0733 0.3216 1.0000
5.750 1.0096 0.01841 0.00980 -0.0715 0.3137 1.0000
6.000 1.0258 0.01871 0.01006 -0.0692 0.3031 1.0000
6.250 1.0436 0.01897 0.01033 -0.0673 0.2938 1.0000
6.500 1.0595 0.01930 0.01062 -0.0650 0.2847 1.0000
6.750 1.0775 0.01956 0.01092 -0.0631 0.2754 1.0000
7.000 1.0929 0.01991 0.01124 -0.0608 0.2666 1.0000
7.250 1.1093 0.02022 0.01156 -0.0587 0.2554 1.0000
7.500 1.1245 0.02059 0.01190 -0.0563 0.2447 1.0000
7.750 1.1385 0.02099 0.01229 -0.0538 0.2346 1.0000
8.000 1.1540 0.02137 0.01266 -0.0516 0.2239 1.0000
8.250 1.1677 0.02181 0.01308 -0.0491 0.2141 1.0000
8.500 1.1809 0.02228 0.01353 -0.0466 0.2046 1.0000
8.750 1.1947 0.02274 0.01401 -0.0442 0.1969 1.0000
9.000 1.2080 0.02324 0.01451 -0.0417 0.1907 1.0000
9.250 1.2205 0.02377 0.01506 -0.0392 0.1842 1.0000
9.500 1.2337 0.02428 0.01561 -0.0368 0.1787 1.0000
9.750 1.2442 0.02491 0.01624 -0.0340 0.1728 1.0000
10.000 1.2573 0.02545 0.01684 -0.0317 0.1678 1.0000
10.250 1.2683 0.02608 0.01750 -0.0291 0.1617 1.0000
10.500 1.2769 0.02683 0.01825 -0.0262 0.1566 1.0000
10.750 1.2898 0.02740 0.01892 -0.0240 0.1515 1.0000
11.000 1.2995 0.02813 0.01970 -0.0215 0.1455 1.0000
11.250 1.3076 0.02896 0.02056 -0.0187 0.1398 1.0000
11.500 1.3181 0.02970 0.02137 -0.0164 0.1330 1.0000
11.750 1.3245 0.03066 0.02236 -0.0136 0.1265 1.0000
12.000 1.3332 0.03153 0.02332 -0.0113 0.1198 1.0000
12.250 1.3381 0.03265 0.02446 -0.0086 0.1137 1.0000
12.500 1.3438 0.03378 0.02563 -0.0061 0.1061 1.0000
12.750 1.3466 0.03512 0.02700 -0.0035 0.1003 1.0000
13.000 1.3504 0.03648 0.02842 -0.0011 0.0963 1.0000
13.250 1.3518 0.03805 0.03004 0.0013 0.0928 1.0000
13.500 1.3519 0.03981 0.03184 0.0036 0.0897 1.0000
13.750 1.3541 0.04149 0.03362 0.0055 0.0875 1.0000
14.000 1.3544 0.04341 0.03563 0.0073 0.0855 1.0000
14.250 1.3542 0.04545 0.03775 0.0089 0.0838 1.0000
14.500 1.3516 0.04783 0.04020 0.0104 0.0821 1.0000
14.750 1.3465 0.05056 0.04299 0.0116 0.0803 1.0000
15.000 1.3451 0.05301 0.04555 0.0126 0.0794 1.0000
15.250 1.3441 0.05553 0.04819 0.0133 0.0771 1.0000
15.500 1.3425 0.05818 0.05095 0.0139 0.0761 1.0000
15.750 1.3388 0.06113 0.05400 0.0143 0.0745 1.0000
16.000 1.3356 0.06409 0.05705 0.0146 0.0738 1.0000
16.250 1.3291 0.06750 0.06053 0.0146 0.0723 1.0000
16.500 1.3239 0.07082 0.06393 0.0146 0.0715 1.0000
16.750 1.3161 0.07452 0.06767 0.0144 0.0700 1.0000
17.000 1.3139 0.07757 0.07083 0.0143 0.0696 1.0000
17.250 1.3100 0.08097 0.07438 0.0138 0.0683 1.0000
17.500 1.3060 0.08443 0.07796 0.0132 0.0672 1.0000
17.750 1.3021 0.08786 0.08151 0.0127 0.0665 1.0000
18.000 1.2979 0.09141 0.08517 0.0120 0.0657 1.0000
18.250 1.2926 0.09514 0.08899 0.0112 0.0647 1.0000
18.500 1.2868 0.09897 0.09291 0.0102 0.0636 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 575 AIRFOIL (goe575-il)