GOE 802 B AIRFOIL (goe802b-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 802 B AIRFOIL (goe802b-il) Reynolds number: 200,000 Max Cl/Cd: 73.36 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe802b-il-200000-n5.txt Download as CSV file: xf-goe802b-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 802 B 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.1589 0.09258 0.08909 -0.0576 0.9616 0.0277
-8.250 -0.1449 0.08950 0.08599 -0.0625 0.9502 0.0281
-8.000 -0.1318 0.08664 0.08311 -0.0681 0.9378 0.0282
-7.750 -0.1174 0.08284 0.07930 -0.0710 0.9305 0.0285
-7.500 -0.1070 0.07972 0.07617 -0.0709 0.9228 0.0289
-7.250 -0.0934 0.07686 0.07330 -0.0729 0.9145 0.0295
-7.000 -0.0775 0.07392 0.07033 -0.0758 0.9073 0.0301
-6.750 -0.0613 0.07104 0.06742 -0.0789 0.8983 0.0309
-6.500 -0.0421 0.06799 0.06431 -0.0825 0.8912 0.0319
-6.250 -0.0173 0.06485 0.06109 -0.0879 0.8830 0.0332
-6.000 0.0201 0.06167 0.05769 -0.0971 0.8747 0.0337
-5.750 0.0306 0.05823 0.05427 -0.0961 0.8691 0.0341
-5.500 0.0473 0.05559 0.05161 -0.0970 0.8610 0.0348
-5.250 0.0697 0.05298 0.04892 -0.0991 0.8541 0.0357
-5.000 0.0947 0.05036 0.04619 -0.1016 0.8473 0.0368
-4.750 0.1230 0.04777 0.04348 -0.1046 0.8392 0.0387
-4.500 0.1640 0.04567 0.04097 -0.1091 0.8326 0.0396
-4.250 0.1833 0.04256 0.03786 -0.1099 0.8249 0.0400
-4.000 0.2044 0.04018 0.03544 -0.1106 0.8167 0.0408
-3.750 0.2296 0.03829 0.03343 -0.1116 0.8094 0.0427
-3.500 0.2655 0.03722 0.03199 -0.1133 0.8000 0.0457
-3.250 0.2919 0.03493 0.02952 -0.1142 0.7918 0.0460
-3.000 0.3150 0.03262 0.02717 -0.1148 0.7805 0.0466
-2.750 0.3410 0.03092 0.02533 -0.1154 0.7703 0.0477
-2.500 0.3683 0.02949 0.02372 -0.1159 0.7586 0.0495
-2.250 0.4005 0.02923 0.02305 -0.1159 0.7479 0.0518
-1.750 0.4524 0.02559 0.01920 -0.1169 0.7252 0.0536
-1.500 0.4810 0.02519 0.01853 -0.1168 0.7137 0.0577
-1.250 0.5094 0.02449 0.01759 -0.1169 0.7012 0.0580
-0.750 0.5666 0.02083 0.01324 -0.1170 0.6769 0.0394
-0.500 0.5937 0.01994 0.01219 -0.1170 0.6631 0.0390
-0.250 0.6209 0.01920 0.01124 -0.1169 0.6500 0.0387
0.000 0.6480 0.01854 0.01036 -0.1168 0.6373 0.0387
0.250 0.6751 0.01792 0.00954 -0.1166 0.6244 0.0393
0.500 0.7016 0.01759 0.00911 -0.1165 0.6115 0.0409
0.750 0.7280 0.01725 0.00862 -0.1163 0.5984 0.0418
1.000 0.7546 0.01679 0.00802 -0.1161 0.5854 0.0417
1.250 0.7809 0.01642 0.00751 -0.1159 0.5736 0.0419
1.500 0.8071 0.01611 0.00708 -0.1157 0.5621 0.0424
1.750 0.8333 0.01583 0.00672 -0.1155 0.5515 0.0434
2.000 0.8591 0.01556 0.00636 -0.1152 0.5410 0.0448
2.250 0.8851 0.01551 0.00629 -0.1150 0.5302 0.0471
2.750 0.9365 0.01554 0.00627 -0.1144 0.5088 0.0538
3.000 0.9617 0.01556 0.00623 -0.1141 0.4991 0.0587
3.250 0.9877 0.01565 0.00633 -0.1139 0.4893 0.0670
3.500 1.0131 0.01599 0.00656 -0.1135 0.4801 0.0791
3.750 1.0390 0.01619 0.00671 -0.1132 0.4695 0.0902
4.000 1.0634 0.01629 0.00674 -0.1128 0.4580 0.0958
4.250 1.0877 0.01629 0.00674 -0.1124 0.4450 0.0980
4.500 1.1119 0.01637 0.00684 -0.1120 0.4325 0.0993
4.750 1.1354 0.01657 0.00699 -0.1114 0.4212 0.1011
5.000 1.1596 0.01664 0.00713 -0.1110 0.4098 0.1020
5.250 1.1830 0.01680 0.00732 -0.1105 0.3992 0.1033
5.750 1.2302 0.01716 0.00775 -0.1096 0.3819 0.1063
6.000 1.2540 0.01731 0.00798 -0.1092 0.3728 0.1086
6.250 1.2762 0.01755 0.00821 -0.1086 0.3632 0.1128
6.500 1.2994 0.01772 0.00840 -0.1081 0.3526 0.1188
6.750 1.3210 0.01803 0.00866 -0.1073 0.3429 0.1228
7.000 1.3432 0.01831 0.00896 -0.1066 0.3328 0.1251
7.250 1.3640 0.01869 0.00933 -0.1056 0.3229 0.1267
7.500 1.3852 0.01904 0.00970 -0.1048 0.3123 0.1284
7.750 1.4054 0.01945 0.01010 -0.1038 0.3028 0.1303
8.000 1.4258 0.01983 0.01054 -0.1028 0.2930 0.1326
8.250 1.4450 0.02027 0.01102 -0.1017 0.2835 0.1367
8.500 1.4648 0.02013 0.01174 -0.1011 0.2712 0.7146
9.000 1.4919 0.02080 0.01277 -0.0966 0.2445 1.0000
9.250 1.5029 0.02154 0.01345 -0.0943 0.2254 1.0000
9.500 1.5119 0.02246 0.01427 -0.0918 0.2050 1.0000
9.750 1.5196 0.02352 0.01525 -0.0893 0.1870 1.0000
10.000 1.5276 0.02463 0.01630 -0.0871 0.1728 1.0000
10.250 1.5354 0.02580 0.01745 -0.0849 0.1623 1.0000
10.500 1.5452 0.02689 0.01855 -0.0830 0.1527 1.0000
10.750 1.5528 0.02816 0.01983 -0.0811 0.1446 1.0000
11.000 1.5629 0.02930 0.02104 -0.0795 0.1368 1.0000
11.250 1.5692 0.03075 0.02250 -0.0777 0.1296 1.0000
11.500 1.5785 0.03201 0.02384 -0.0762 0.1225 1.0000
11.750 1.5840 0.03362 0.02546 -0.0746 0.1143 1.0000
12.000 1.5902 0.03522 0.02710 -0.0731 0.1058 1.0000
12.250 1.5933 0.03714 0.02903 -0.0716 0.0973 1.0000
12.500 1.5989 0.03890 0.03087 -0.0703 0.0901 1.0000
12.750 1.6008 0.04106 0.03305 -0.0691 0.0824 1.0000
13.000 1.6022 0.04332 0.03535 -0.0679 0.0752 1.0000
13.250 1.6011 0.04592 0.03797 -0.0668 0.0688 1.0000
13.500 1.6002 0.04859 0.04069 -0.0659 0.0635 1.0000
13.750 1.5965 0.05165 0.04378 -0.0651 0.0586 1.0000
14.000 1.5959 0.05448 0.04670 -0.0645 0.0540 1.0000
14.250 1.5929 0.05770 0.05001 -0.0641 0.0490 1.0000
14.500 1.5898 0.06104 0.05343 -0.0639 0.0433 1.0000
14.750 1.5811 0.06517 0.05759 -0.0638 0.0270 1.0000
15.000 1.5660 0.07032 0.06271 -0.0641 0.0221 1.0000
15.250 1.5548 0.07511 0.06757 -0.0646 0.0202 1.0000
15.500 1.5451 0.07979 0.07236 -0.0652 0.0192 1.0000
15.750 1.5356 0.08455 0.07726 -0.0660 0.0184 1.0000
16.000 1.5260 0.08944 0.08229 -0.0669 0.0178 1.0000
16.250 1.5159 0.09447 0.08746 -0.0680 0.0173 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 802 B AIRFOIL (goe802b-il)