GOE 360 AIRFOIL (goe360-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 360 AIRFOIL (goe360-il) Reynolds number: 200,000 Max Cl/Cd: 76.19 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe360-il-200000-n5.txt Download as CSV file: xf-goe360-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 360 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 -0.2816 0.12285 0.11940 -0.0262 1.0000 0.0215
-10.250 -0.2782 0.12034 0.11692 -0.0288 1.0000 0.0216
-10.000 -0.2741 0.11752 0.11414 -0.0310 1.0000 0.0217
-9.750 -0.2618 0.11270 0.10934 -0.0301 1.0000 0.0219
-9.500 -0.2536 0.10943 0.10610 -0.0305 1.0000 0.0222
-9.250 -0.2466 0.10654 0.10323 -0.0314 1.0000 0.0225
-9.000 -0.2379 0.10356 0.10028 -0.0331 0.9916 0.0230
-8.750 -0.2239 0.10010 0.09681 -0.0365 0.9729 0.0236
-8.500 -0.2098 0.09670 0.09340 -0.0400 0.9569 0.0248
-8.250 -0.1973 0.09391 0.09058 -0.0456 0.9393 0.0264
-8.000 -0.1893 0.09122 0.08788 -0.0505 0.9228 0.0266
-7.750 -0.1796 0.08815 0.08478 -0.0553 0.9094 0.0267
-7.500 -0.1662 0.08470 0.08127 -0.0604 0.8977 0.0268
-7.000 -0.1431 0.07703 0.07353 -0.0643 0.8761 0.0271
-6.750 -0.1355 0.07407 0.07055 -0.0615 0.8655 0.0276
-6.500 -0.1231 0.07136 0.06779 -0.0622 0.8538 0.0285
-6.250 -0.1075 0.06844 0.06481 -0.0647 0.8418 0.0295
-6.000 -0.0890 0.06528 0.06159 -0.0681 0.8290 0.0305
-5.750 -0.0666 0.06204 0.05824 -0.0725 0.8160 0.0321
-5.500 -0.0242 0.05881 0.05471 -0.0825 0.8025 0.0338
-5.250 0.0044 0.05574 0.05140 -0.0862 0.7890 0.0340
-5.000 0.0210 0.05125 0.04682 -0.0876 0.7756 0.0344
-4.750 0.0367 0.04815 0.04366 -0.0878 0.7611 0.0349
-4.500 0.0564 0.04562 0.04103 -0.0884 0.7460 0.0356
-4.250 0.0791 0.04325 0.03851 -0.0895 0.7308 0.0366
-4.000 0.1040 0.04097 0.03604 -0.0908 0.7162 0.0380
-3.750 0.1434 0.03967 0.03424 -0.0933 0.7024 0.0432
-3.250 0.2000 0.03592 0.02980 -0.0948 0.6778 0.0436
-3.000 0.2207 0.03196 0.02580 -0.0958 0.6674 0.0446
-2.750 0.2446 0.03008 0.02379 -0.0962 0.6572 0.0455
-2.500 0.2701 0.02857 0.02213 -0.0966 0.6471 0.0470
-2.250 0.3033 0.02886 0.02190 -0.0965 0.6377 0.0540
-2.000 0.3316 0.02794 0.02064 -0.0965 0.6286 0.0543
-1.750 0.3557 0.02463 0.01731 -0.0973 0.6201 0.0557
-1.250 0.4129 0.02146 0.01361 -0.0971 0.6035 0.0416
-1.000 0.4406 0.02032 0.01222 -0.0971 0.5952 0.0404
-0.750 0.4688 0.01942 0.01110 -0.0970 0.5860 0.0407
-0.500 0.4968 0.01873 0.01017 -0.0968 0.5777 0.0416
-0.250 0.5247 0.01792 0.00919 -0.0967 0.5685 0.0409
0.000 0.5524 0.01723 0.00833 -0.0965 0.5594 0.0402
0.250 0.5799 0.01664 0.00758 -0.0963 0.5497 0.0398
0.500 0.6074 0.01612 0.00694 -0.0961 0.5396 0.0397
0.750 0.6346 0.01569 0.00639 -0.0959 0.5299 0.0398
1.000 0.6617 0.01533 0.00593 -0.0957 0.5195 0.0402
1.250 0.6886 0.01502 0.00556 -0.0954 0.5085 0.0407
1.500 0.7153 0.01489 0.00534 -0.0952 0.4975 0.0425
1.750 0.7418 0.01482 0.00518 -0.0949 0.4866 0.0436
2.000 0.7682 0.01455 0.00491 -0.0947 0.4758 0.0440
2.250 0.7945 0.01436 0.00472 -0.0945 0.4660 0.0447
2.500 0.8206 0.01429 0.00461 -0.0943 0.4563 0.0457
2.750 0.8470 0.01428 0.00457 -0.0942 0.4463 0.0472
3.000 0.8732 0.01435 0.00457 -0.0940 0.4373 0.0492
3.250 0.8994 0.01445 0.00461 -0.0938 0.4287 0.0521
3.500 0.9257 0.01455 0.00468 -0.0936 0.4211 0.0589
3.750 0.9517 0.01470 0.00479 -0.0934 0.4136 0.0680
4.000 0.9755 0.01328 0.00503 -0.0927 0.4069 1.0000
4.250 1.0012 0.01353 0.00519 -0.0925 0.3997 1.0000
4.500 1.0268 0.01381 0.00540 -0.0922 0.3934 1.0000
4.750 1.0527 0.01406 0.00562 -0.0920 0.3871 1.0000
5.000 1.0780 0.01435 0.00586 -0.0917 0.3817 1.0000
5.250 1.1037 0.01462 0.00612 -0.0915 0.3765 1.0000
5.500 1.1291 0.01490 0.00641 -0.0912 0.3709 1.0000
5.750 1.1540 0.01523 0.00670 -0.0909 0.3659 1.0000
6.000 1.1793 0.01551 0.00702 -0.0907 0.3609 1.0000
6.250 1.2045 0.01581 0.00736 -0.0904 0.3562 1.0000
6.500 1.2292 0.01615 0.00770 -0.0901 0.3521 1.0000
6.750 1.2538 0.01650 0.00807 -0.0898 0.3483 1.0000
7.000 1.2787 0.01680 0.00848 -0.0895 0.3438 1.0000
7.250 1.3030 0.01714 0.00887 -0.0891 0.3392 1.0000
7.500 1.3263 0.01754 0.00925 -0.0886 0.3339 1.0000
7.750 1.3497 0.01783 0.00966 -0.0881 0.3260 1.0000
8.000 1.3714 0.01821 0.01006 -0.0874 0.3176 1.0000
8.250 1.3936 0.01854 0.01049 -0.0868 0.3085 1.0000
8.500 1.4148 0.01893 0.01093 -0.0860 0.2995 1.0000
8.750 1.4342 0.01936 0.01138 -0.0850 0.2872 1.0000
9.000 1.4529 0.01982 0.01190 -0.0839 0.2739 1.0000
9.250 1.4708 0.02033 0.01245 -0.0827 0.2561 1.0000
9.500 1.4848 0.02105 0.01312 -0.0810 0.2315 1.0000
9.750 1.4890 0.02228 0.01413 -0.0781 0.1871 1.0000
10.000 1.4670 0.02531 0.01655 -0.0726 0.1067 1.0000
10.250 1.4505 0.02840 0.01935 -0.0685 0.0574 1.0000
10.500 1.4391 0.03134 0.02214 -0.0657 0.0317 1.0000
10.750 1.4385 0.03358 0.02447 -0.0639 0.0265 1.0000
11.000 1.4401 0.03575 0.02676 -0.0626 0.0240 1.0000
11.250 1.4391 0.03828 0.02941 -0.0614 0.0222 1.0000
11.500 1.4365 0.04114 0.03240 -0.0605 0.0208 1.0000
11.750 1.4365 0.04385 0.03527 -0.0599 0.0198 1.0000
12.000 1.4343 0.04693 0.03850 -0.0595 0.0189 1.0000
12.250 1.4304 0.05034 0.04206 -0.0593 0.0182 1.0000
12.500 1.4246 0.05406 0.04594 -0.0593 0.0177 1.0000
12.750 1.4171 0.05809 0.05012 -0.0594 0.0173 1.0000
13.000 1.4080 0.06244 0.05463 -0.0598 0.0169 1.0000
13.250 1.3974 0.06714 0.05947 -0.0603 0.0166 1.0000
13.500 1.3858 0.07212 0.06458 -0.0611 0.0163 1.0000
13.750 1.3731 0.07734 0.06994 -0.0620 0.0160 1.0000
14.000 1.3615 0.08251 0.07523 -0.0629 0.0158 1.0000
14.250 1.3547 0.08705 0.07989 -0.0638 0.0155 1.0000
14.500 1.3484 0.09154 0.08451 -0.0647 0.0152 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 360 AIRFOIL (goe360-il)