GOE 400 AIRFOIL (goe400-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 400 AIRFOIL (goe400-il) Reynolds number: 1,000,000 Max Cl/Cd: 101.4 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe400-il-1000000-n5.txt Download as CSV file: xf-goe400-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 400 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.4986 0.13165 0.13004 0.0168 1.0000 0.0053
-11.000 -0.5007 0.12560 0.12400 0.0146 1.0000 0.0058
-10.750 -0.4995 0.12083 0.11923 0.0127 1.0000 0.0061
-10.500 -0.4924 0.11786 0.11627 0.0113 1.0000 0.0063
-10.250 -0.4853 0.11485 0.11325 0.0099 1.0000 0.0064
-7.750 -0.4270 0.01843 0.01530 -0.0828 0.9105 0.0182
-7.500 -0.4030 0.01740 0.01401 -0.0824 0.8769 0.0190
-7.250 -0.3782 0.01661 0.01297 -0.0821 0.8418 0.0195
-7.000 -0.3524 0.01600 0.01211 -0.0818 0.8084 0.0198
-6.750 -0.3259 0.01568 0.01159 -0.0816 0.7802 0.0201
-6.500 -0.2999 0.01439 0.00999 -0.0818 0.7581 0.0207
-6.250 -0.2730 0.01394 0.00939 -0.0817 0.7395 0.0213
-5.750 -0.2182 0.01367 0.00897 -0.0814 0.7106 0.0223
-5.500 -0.1906 0.01341 0.00862 -0.0813 0.6988 0.0228
-5.250 -0.1632 0.01295 0.00801 -0.0812 0.6879 0.0233
-5.000 -0.1357 0.01249 0.00741 -0.0811 0.6776 0.0238
-4.750 -0.1080 0.01206 0.00685 -0.0811 0.6677 0.0244
-4.500 -0.0803 0.01154 0.00617 -0.0810 0.6589 0.0247
-4.250 -0.0526 0.01108 0.00556 -0.0809 0.6496 0.0250
-4.000 -0.0248 0.01064 0.00500 -0.0808 0.6411 0.0252
-3.750 0.0030 0.01027 0.00450 -0.0807 0.6324 0.0255
-3.500 0.0309 0.00992 0.00404 -0.0806 0.6239 0.0257
-3.250 0.0588 0.00961 0.00363 -0.0804 0.6157 0.0259
-3.000 0.0868 0.00934 0.00326 -0.0803 0.6071 0.0261
-2.750 0.1148 0.00914 0.00299 -0.0802 0.5993 0.0264
-2.500 0.1428 0.00897 0.00276 -0.0800 0.5912 0.0266
-2.250 0.1708 0.00878 0.00251 -0.0799 0.5838 0.0268
-2.000 0.1988 0.00862 0.00229 -0.0797 0.5758 0.0270
-1.750 0.2269 0.00847 0.00208 -0.0796 0.5684 0.0271
-1.500 0.2549 0.00818 0.00170 -0.0795 0.5604 0.0283
-1.250 0.2829 0.00803 0.00151 -0.0793 0.5529 0.0293
-1.000 0.3110 0.00794 0.00136 -0.0792 0.5446 0.0302
-0.750 0.3390 0.00786 0.00126 -0.0791 0.5370 0.0310
-0.500 0.3671 0.00781 0.00117 -0.0789 0.5283 0.0317
-0.250 0.3951 0.00777 0.00110 -0.0788 0.5191 0.0325
0.000 0.4230 0.00776 0.00104 -0.0787 0.5063 0.0333
0.250 0.4509 0.00777 0.00100 -0.0785 0.4921 0.0343
0.500 0.4787 0.00778 0.00097 -0.0784 0.4786 0.0356
1.000 0.5342 0.00776 0.00099 -0.0782 0.4511 0.0791
1.250 0.5620 0.00784 0.00106 -0.0780 0.4389 0.0881
1.500 0.5897 0.00794 0.00111 -0.0779 0.4235 0.0920
1.750 0.6173 0.00808 0.00118 -0.0778 0.4051 0.0941
2.000 0.6448 0.00818 0.00123 -0.0776 0.3868 0.0964
2.250 0.6723 0.00827 0.00128 -0.0775 0.3725 0.0983
2.500 0.6999 0.00836 0.00135 -0.0774 0.3600 0.1000
2.750 0.7274 0.00846 0.00142 -0.0772 0.3474 0.1015
3.000 0.7549 0.00858 0.00151 -0.0771 0.3347 0.1039
3.250 0.7823 0.00871 0.00162 -0.0770 0.3214 0.1062
3.500 0.8095 0.00885 0.00172 -0.0769 0.3069 0.1068
3.750 0.8367 0.00900 0.00183 -0.0767 0.2922 0.1072
4.000 0.8639 0.00914 0.00195 -0.0766 0.2783 0.1081
4.250 0.8910 0.00930 0.00207 -0.0764 0.2640 0.1095
4.500 0.9176 0.00953 0.00224 -0.0762 0.2432 0.1110
4.750 0.9443 0.00975 0.00241 -0.0761 0.2248 0.1127
5.000 0.9704 0.01009 0.00265 -0.0758 0.1973 0.1145
5.250 0.9952 0.01065 0.00301 -0.0755 0.1526 0.1166
5.500 1.0188 0.01142 0.00348 -0.0751 0.0957 0.1191
6.000 1.0712 0.01191 0.00402 -0.0747 0.0817 0.1520
6.500 1.1215 0.01106 0.00470 -0.0741 0.0723 1.0000
6.750 1.1472 0.01139 0.00500 -0.0738 0.0654 1.0000
7.000 1.1714 0.01196 0.00540 -0.0734 0.0400 1.0000
7.250 1.1946 0.01269 0.00601 -0.0729 0.0186 1.0000
7.500 1.2193 0.01312 0.00643 -0.0725 0.0145 1.0000
7.750 1.2444 0.01349 0.00682 -0.0721 0.0126 1.0000
8.000 1.2689 0.01391 0.00726 -0.0717 0.0110 1.0000
8.250 1.2930 0.01440 0.00777 -0.0712 0.0095 1.0000
8.500 1.3174 0.01480 0.00820 -0.0708 0.0090 1.0000
8.750 1.3415 0.01523 0.00867 -0.0704 0.0084 1.0000
9.000 1.3650 0.01571 0.00917 -0.0699 0.0077 1.0000
9.250 1.3879 0.01626 0.00977 -0.0693 0.0071 1.0000
9.500 1.4103 0.01685 0.01040 -0.0686 0.0065 1.0000
9.750 1.4330 0.01735 0.01094 -0.0681 0.0062 1.0000
10.000 1.4551 0.01790 0.01154 -0.0674 0.0059 1.0000
10.250 1.4767 0.01849 0.01218 -0.0667 0.0056 1.0000
10.500 1.4976 0.01912 0.01285 -0.0660 0.0053 1.0000
10.750 1.5175 0.01982 0.01361 -0.0651 0.0050 1.0000
11.000 1.5357 0.02067 0.01453 -0.0640 0.0047 1.0000
11.250 1.5529 0.02156 0.01549 -0.0629 0.0045 1.0000
11.500 1.5705 0.02233 0.01634 -0.0618 0.0044 1.0000
11.750 1.5865 0.02319 0.01727 -0.0605 0.0043 1.0000
12.000 1.6011 0.02411 0.01827 -0.0591 0.0042 1.0000
12.250 1.6130 0.02506 0.01930 -0.0573 0.0040 1.0000
12.500 1.6218 0.02610 0.02043 -0.0553 0.0039 1.0000
12.750 1.6295 0.02727 0.02169 -0.0534 0.0038 1.0000
13.000 1.6364 0.02861 0.02311 -0.0518 0.0036 1.0000
13.250 1.6427 0.03015 0.02473 -0.0505 0.0035 1.0000
13.500 1.6477 0.03196 0.02663 -0.0496 0.0034 1.0000
13.750 1.6516 0.03407 0.02883 -0.0489 0.0034 1.0000
14.000 1.6534 0.03656 0.03142 -0.0486 0.0033 1.0000
14.250 1.6533 0.03944 0.03440 -0.0486 0.0032 1.0000
14.500 1.6503 0.04278 0.03786 -0.0488 0.0031 1.0000
14.750 1.6437 0.04667 0.04188 -0.0493 0.0031 1.0000
15.000 1.6353 0.05089 0.04622 -0.0499 0.0030 1.0000
15.250 1.6271 0.05515 0.05061 -0.0507 0.0030 1.0000
15.500 1.6193 0.05941 0.05499 -0.0515 0.0030 1.0000
15.750 1.6100 0.06397 0.05968 -0.0526 0.0030 1.0000
16.000 1.5988 0.06889 0.06471 -0.0538 0.0029 1.0000
16.250 1.5872 0.07408 0.07003 -0.0553 0.0029 1.0000
16.500 1.5747 0.07960 0.07568 -0.0570 0.0029 1.0000
16.750 1.5612 0.08541 0.08162 -0.0590 0.0029 1.0000
17.000 1.5466 0.09158 0.08791 -0.0612 0.0029 1.0000
17.250 1.5317 0.09792 0.09437 -0.0635 0.0029 1.0000
17.500 1.5163 0.10443 0.10100 -0.0660 0.0029 1.0000
17.750 1.5001 0.11125 0.10794 -0.0688 0.0029 1.0000
18.000 1.4844 0.11810 0.11490 -0.0717 0.0029 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 400 AIRFOIL (goe400-il)