GOE 362 AIRFOIL (goe362-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 362 AIRFOIL (goe362-il) Reynolds number: 200,000 Max Cl/Cd: 81.93 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe362-il-200000-n5.txt Download as CSV file: xf-goe362-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 362 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
------ -------- --------- --------- -------- -------- --------
-7.500 -0.2776 0.09241 0.08911 -0.0255 1.0000 0.0214
-7.250 -0.2814 0.09063 0.08740 -0.0242 1.0000 0.0218
-7.000 -0.2904 0.08940 0.08626 -0.0219 1.0000 0.0225
-6.750 -0.2700 0.08602 0.08290 -0.0295 0.9930 0.0238
-6.500 -0.2399 0.08165 0.07852 -0.0394 0.9855 0.0241
-6.250 -0.2103 0.07707 0.07392 -0.0475 0.9785 0.0243
-6.000 -0.1811 0.07248 0.06930 -0.0547 0.9712 0.0244
-5.750 -0.1474 0.06753 0.06431 -0.0625 0.9659 0.0244
-5.500 -0.1376 0.06342 0.06021 -0.0611 0.9580 0.0205
-5.250 -0.1019 0.05863 0.05535 -0.0694 0.9496 0.0207
-5.000 -0.0653 0.05365 0.05030 -0.0776 0.9397 0.0200
-4.500 0.0223 0.03968 0.03591 -0.0970 0.9166 0.0155
-4.250 0.0613 0.03368 0.02963 -0.1035 0.9050 0.0154
-4.000 0.1022 0.02570 0.02108 -0.1097 0.8941 0.0154
-3.750 0.1368 0.01976 0.01422 -0.1128 0.8822 0.0163
-3.500 0.1656 0.01839 0.01259 -0.1134 0.8699 0.0181
-3.250 0.1952 0.01637 0.01005 -0.1137 0.8578 0.0194
-3.000 0.2242 0.01486 0.00810 -0.1136 0.8454 0.0210
-2.750 0.2526 0.01397 0.00688 -0.1133 0.8325 0.0238
-2.500 0.2803 0.01304 0.00575 -0.1130 0.8194 0.0270
-2.250 0.3078 0.01250 0.00505 -0.1126 0.8059 0.0312
-2.000 0.3351 0.01201 0.00442 -0.1122 0.7923 0.0379
-1.750 0.3622 0.01170 0.00397 -0.1118 0.7784 0.0455
-1.500 0.3892 0.01142 0.00361 -0.1113 0.7643 0.0566
-1.250 0.4161 0.01124 0.00338 -0.1109 0.7500 0.0713
-0.750 0.4697 0.01101 0.00300 -0.1100 0.7211 0.1048
-0.500 0.4962 0.01091 0.00285 -0.1096 0.7063 0.1234
-0.250 0.5228 0.01084 0.00274 -0.1091 0.6918 0.1409
0.000 0.5492 0.01079 0.00270 -0.1087 0.6772 0.1725
0.250 0.5756 0.01074 0.00266 -0.1083 0.6625 0.2162
0.500 0.6017 0.01064 0.00265 -0.1078 0.6481 0.2715
0.750 0.6272 0.01036 0.00270 -0.1074 0.6335 0.4252
1.000 0.6583 0.00938 0.00265 -0.1078 0.6185 1.0000
1.250 0.6847 0.00954 0.00266 -0.1073 0.6034 1.0000
1.500 0.7109 0.00971 0.00269 -0.1069 0.5878 1.0000
1.750 0.7372 0.00988 0.00274 -0.1064 0.5717 1.0000
2.000 0.7633 0.01005 0.00281 -0.1059 0.5547 1.0000
2.250 0.7893 0.01024 0.00290 -0.1055 0.5374 1.0000
2.500 0.8151 0.01044 0.00300 -0.1050 0.5201 1.0000
2.750 0.8408 0.01066 0.00311 -0.1044 0.5034 1.0000
3.000 0.8664 0.01089 0.00326 -0.1039 0.4880 1.0000
3.250 0.8919 0.01114 0.00342 -0.1034 0.4738 1.0000
3.500 0.9175 0.01139 0.00361 -0.1030 0.4610 1.0000
3.750 0.9430 0.01165 0.00381 -0.1025 0.4492 1.0000
4.000 0.9684 0.01192 0.00406 -0.1020 0.4385 1.0000
4.250 0.9939 0.01220 0.00430 -0.1015 0.4280 1.0000
4.500 1.0194 0.01246 0.00457 -0.1011 0.4182 1.0000
4.750 1.0446 0.01275 0.00487 -0.1006 0.4082 1.0000
5.000 1.0694 0.01307 0.00516 -0.1000 0.3959 1.0000
5.250 1.0939 0.01336 0.00545 -0.0994 0.3805 1.0000
5.500 1.1184 0.01366 0.00577 -0.0989 0.3644 1.0000
5.750 1.1428 0.01396 0.00608 -0.0983 0.3485 1.0000
6.000 1.1668 0.01427 0.00641 -0.0976 0.3310 1.0000
6.250 1.1907 0.01461 0.00676 -0.0970 0.3134 1.0000
6.500 1.2141 0.01498 0.00717 -0.0963 0.2958 1.0000
6.750 1.2372 0.01539 0.00759 -0.0956 0.2780 1.0000
7.000 1.2599 0.01584 0.00805 -0.0948 0.2572 1.0000
7.250 1.2813 0.01639 0.00857 -0.0938 0.2327 1.0000
7.500 1.3014 0.01708 0.00918 -0.0928 0.2027 1.0000
7.750 1.3199 0.01793 0.00994 -0.0915 0.1680 1.0000
8.000 1.3361 0.01902 0.01083 -0.0901 0.1278 1.0000
8.500 1.3670 0.02125 0.01280 -0.0870 0.0800 1.0000
8.750 1.3837 0.02216 0.01372 -0.0856 0.0688 1.0000
9.000 1.4004 0.02303 0.01463 -0.0842 0.0593 1.0000
9.250 1.4168 0.02386 0.01552 -0.0827 0.0491 1.0000
9.500 1.4302 0.02492 0.01661 -0.0809 0.0358 1.0000
10.000 1.4445 0.02767 0.01930 -0.0756 0.0158 1.0000
10.250 1.4482 0.02914 0.02087 -0.0725 0.0137 1.0000
10.500 1.4492 0.03085 0.02274 -0.0694 0.0122 1.0000
10.750 1.4531 0.03241 0.02448 -0.0669 0.0114 1.0000
11.000 1.4555 0.03416 0.02640 -0.0646 0.0106 1.0000
11.250 1.4563 0.03618 0.02858 -0.0626 0.0099 1.0000
11.500 1.4553 0.03847 0.03103 -0.0608 0.0093 1.0000
11.750 1.4521 0.04116 0.03388 -0.0593 0.0089 1.0000
12.000 1.4456 0.04439 0.03727 -0.0582 0.0086 1.0000
12.250 1.4378 0.04800 0.04105 -0.0575 0.0084 1.0000
12.500 1.4335 0.05143 0.04464 -0.0572 0.0082 1.0000
12.750 1.4292 0.05502 0.04842 -0.0572 0.0080 1.0000
13.000 1.4244 0.05884 0.05241 -0.0574 0.0079 1.0000
13.250 1.4188 0.06289 0.05663 -0.0579 0.0077 1.0000
13.500 1.4128 0.06714 0.06106 -0.0586 0.0076 1.0000
13.750 1.4064 0.07158 0.06567 -0.0595 0.0074 1.0000
14.000 1.3997 0.07622 0.07048 -0.0607 0.0072 1.0000
14.250 1.3928 0.08107 0.07555 -0.0622 0.0070 1.0000
14.500 1.3857 0.08612 0.08076 -0.0639 0.0068 1.0000
14.750 1.3781 0.09139 0.08619 -0.0658 0.0067 1.0000
15.000 1.3703 0.09690 0.09185 -0.0681 0.0065 1.0000
15.250 1.3620 0.10261 0.09771 -0.0706 0.0064 1.0000
15.500 1.3534 0.10851 0.10376 -0.0732 0.0063 1.0000
15.750 1.3448 0.11457 0.10995 -0.0762 0.0062 1.0000
16.000 1.3360 0.12077 0.11629 -0.0792 0.0061 1.0000
16.250 1.3273 0.12709 0.12274 -0.0825 0.0060 1.0000
16.500 1.3184 0.13357 0.12935 -0.0859 0.0059 1.0000
16.750 1.3095 0.14026 0.13618 -0.0896 0.0059 1.0000
17.000 1.3005 0.14721 0.14327 -0.0936 0.0058 1.0000
17.250 1.2916 0.15440 0.15062 -0.0979 0.0058 1.0000
17.500 1.2824 0.16193 0.15829 -0.1025 0.0058 1.0000
17.750 1.2729 0.16984 0.16634 -0.1074 0.0058 1.0000
18.000 1.2630 0.17832 0.17498 -0.1128 0.0059 1.0000
18.250 1.2513 0.18798 0.18479 -0.1190 0.0059 1.0000
18.500 1.2129 0.21029 0.20733 -0.1321 0.0066 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 362 AIRFOIL (goe362-il)