GOE 364 AIRFOIL (goe364-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 364 AIRFOIL (goe364-il) Reynolds number: 200,000 Max Cl/Cd: 76.37 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe364-il-200000.txt Download as CSV file: xf-goe364-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 364 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.0891 0.09064 0.08761 -0.0518 0.9775 0.0356
-8.750 -0.0728 0.08617 0.08314 -0.0562 0.9729 0.0369
-8.500 -0.0627 0.08227 0.07923 -0.0624 0.9658 0.0381
-8.250 -0.0558 0.07830 0.07525 -0.0720 0.9580 0.0385
-8.000 -0.0280 0.07237 0.06932 -0.0713 0.9563 0.0392
-7.750 -0.0099 0.06857 0.06551 -0.0729 0.9462 0.0403
-7.500 0.0078 0.06463 0.06156 -0.0766 0.9373 0.0423
-7.250 -0.0804 0.07796 0.07479 -0.0731 0.9480 0.0394
-7.000 -0.0523 0.07419 0.07101 -0.0771 0.9426 0.0410
-6.750 -0.0289 0.07050 0.06728 -0.0828 0.9297 0.0429
-6.500 0.0096 0.06563 0.06214 -0.1023 0.9107 0.0455
-6.250 0.0218 0.06161 0.05818 -0.1004 0.8992 0.0461
-6.000 0.0402 0.05883 0.05539 -0.1000 0.8883 0.0472
-5.750 0.0619 0.05601 0.05248 -0.1023 0.8767 0.0489
-5.500 0.0853 0.05301 0.04933 -0.1063 0.8632 0.0523
-5.250 0.1071 0.04918 0.04525 -0.1106 0.8496 0.0550
-5.000 0.1230 0.04709 0.04314 -0.1098 0.8381 0.0565
-4.750 0.1439 0.04490 0.04082 -0.1106 0.8277 0.0590
-4.500 0.1720 0.04181 0.03725 -0.1141 0.8163 0.0649
-4.250 0.1883 0.03960 0.03507 -0.1133 0.8061 0.0662
-4.000 0.2092 0.03785 0.03322 -0.1132 0.7974 0.0689
-3.750 0.2349 0.03579 0.03075 -0.1140 0.7870 0.0771
-3.500 0.2542 0.03389 0.02886 -0.1135 0.7786 0.0795
-3.250 0.2764 0.03252 0.02737 -0.1131 0.7685 0.0851
-3.000 0.3005 0.03058 0.02510 -0.1130 0.7597 0.0920
-2.750 0.3222 0.02918 0.02365 -0.1124 0.7503 0.0963
-2.500 0.3468 0.02781 0.02196 -0.1119 0.7413 0.1067
-2.250 0.3697 0.02657 0.02068 -0.1114 0.7325 0.1134
-2.000 0.3927 0.02529 0.01922 -0.1107 0.7231 0.1247
-1.750 0.4177 0.02424 0.01793 -0.1102 0.7146 0.1387
-1.500 0.4406 0.02332 0.01688 -0.1093 0.7043 0.1542
-1.250 0.4647 0.02228 0.01571 -0.1087 0.6954 0.1713
-1.000 0.4873 0.02128 0.01468 -0.1078 0.6848 0.1900
-0.750 0.5112 0.02051 0.01382 -0.1070 0.6751 0.2122
-0.500 0.5551 0.01750 0.00943 -0.1048 0.6670 0.0736
-0.250 0.5808 0.01665 0.00844 -0.1038 0.6559 0.0710
0.000 0.6073 0.01601 0.00759 -0.1030 0.6454 0.0694
0.250 0.6325 0.01550 0.00699 -0.1020 0.6335 0.0691
0.500 0.6573 0.01517 0.00661 -0.1011 0.6215 0.0708
0.750 0.6825 0.01489 0.00626 -0.1003 0.6106 0.0729
1.000 0.7073 0.01461 0.00591 -0.0993 0.5995 0.0740
1.250 0.7311 0.01441 0.00570 -0.0983 0.5878 0.0755
1.500 0.7555 0.01420 0.00546 -0.0974 0.5780 0.0789
1.750 0.7803 0.01413 0.00534 -0.0966 0.5684 0.0854
2.000 0.8049 0.01408 0.00530 -0.0957 0.5589 0.1027
2.250 0.8699 0.01237 0.00532 -0.1038 0.5487 1.0000
2.500 0.8937 0.01254 0.00540 -0.1029 0.5392 1.0000
2.750 0.9182 0.01275 0.00546 -0.1021 0.5311 1.0000
3.000 0.9420 0.01293 0.00556 -0.1012 0.5220 1.0000
3.250 0.9662 0.01314 0.00568 -0.1004 0.5138 1.0000
3.500 0.9901 0.01335 0.00581 -0.0995 0.5055 1.0000
3.750 1.0142 0.01358 0.00598 -0.0988 0.4977 1.0000
4.000 1.0380 0.01381 0.00614 -0.0979 0.4897 1.0000
4.250 1.0618 0.01406 0.00634 -0.0971 0.4817 1.0000
4.500 1.0852 0.01430 0.00653 -0.0962 0.4732 1.0000
4.750 1.1083 0.01456 0.00676 -0.0953 0.4646 1.0000
5.000 1.1316 0.01483 0.00695 -0.0945 0.4561 1.0000
5.250 1.1540 0.01511 0.00724 -0.0934 0.4472 1.0000
5.500 1.1775 0.01543 0.00746 -0.0927 0.4387 1.0000
5.750 1.1983 0.01569 0.00777 -0.0914 0.4287 1.0000
6.000 1.2204 0.01602 0.00804 -0.0903 0.4192 1.0000
6.250 1.2409 0.01631 0.00834 -0.0890 0.4088 1.0000
6.500 1.2611 0.01663 0.00867 -0.0877 0.3981 1.0000
6.750 1.2816 0.01699 0.00896 -0.0864 0.3877 1.0000
7.000 1.3000 0.01729 0.00928 -0.0848 0.3762 1.0000
7.250 1.3182 0.01763 0.00966 -0.0831 0.3646 1.0000
7.500 1.3362 0.01800 0.00999 -0.0815 0.3538 1.0000
7.750 1.3533 0.01837 0.01031 -0.0797 0.3429 1.0000
8.000 1.3693 0.01872 0.01072 -0.0778 0.3317 1.0000
8.250 1.3858 0.01916 0.01114 -0.0759 0.3221 1.0000
8.500 1.4017 0.01961 0.01156 -0.0741 0.3135 1.0000
8.750 1.4176 0.02008 0.01208 -0.0722 0.3053 1.0000
9.000 1.4323 0.02059 0.01254 -0.0702 0.2982 1.0000
9.250 1.4464 0.02111 0.01312 -0.0681 0.2910 1.0000
9.500 1.4602 0.02166 0.01366 -0.0660 0.2842 1.0000
9.750 1.4748 0.02227 0.01429 -0.0641 0.2779 1.0000
10.000 1.4885 0.02287 0.01494 -0.0622 0.2718 1.0000
10.250 1.5057 0.02359 0.01559 -0.0609 0.2661 1.0000
10.500 1.5169 0.02424 0.01638 -0.0587 0.2603 1.0000
10.750 1.5303 0.02495 0.01710 -0.0569 0.2545 1.0000
11.000 1.5436 0.02574 0.01793 -0.0552 0.2486 1.0000
11.250 1.5519 0.02652 0.01881 -0.0529 0.2422 1.0000
11.500 1.5649 0.02740 0.01962 -0.0513 0.2357 1.0000
11.750 1.5695 0.02832 0.02072 -0.0488 0.2295 1.0000
12.000 1.5785 0.02928 0.02166 -0.0469 0.2233 1.0000
12.250 1.5845 0.03039 0.02289 -0.0449 0.2174 1.0000
12.500 1.5893 0.03155 0.02415 -0.0429 0.2114 1.0000
12.750 1.5968 0.03276 0.02534 -0.0413 0.2053 1.0000
13.000 1.5981 0.03421 0.02698 -0.0394 0.1995 1.0000
13.250 1.6025 0.03563 0.02843 -0.0379 0.1939 1.0000
13.500 1.6052 0.03726 0.03018 -0.0365 0.1883 1.0000
13.750 1.6062 0.03905 0.03211 -0.0352 0.1826 1.0000
14.000 1.6104 0.04071 0.03373 -0.0341 0.1776 1.0000
14.250 1.6093 0.04286 0.03611 -0.0331 0.1720 1.0000
14.500 1.6088 0.04506 0.03839 -0.0323 0.1666 1.0000
14.750 1.6088 0.04730 0.04069 -0.0316 0.1618 1.0000
15.000 1.6075 0.04982 0.04339 -0.0311 0.1563 1.0000
15.250 1.6060 0.05240 0.04601 -0.0307 0.1518 1.0000
15.500 1.6045 0.05513 0.04887 -0.0305 0.1470 1.0000
15.750 1.6021 0.05806 0.05194 -0.0305 0.1422 1.0000
16.000 1.5967 0.06135 0.05522 -0.0306 0.1375 1.0000
16.250 1.5929 0.06473 0.05880 -0.0310 0.1323 1.0000
16.500 1.5866 0.06847 0.06262 -0.0315 0.1274 1.0000
16.750 1.5805 0.07222 0.06644 -0.0321 0.1233 1.0000
17.000 1.5736 0.07632 0.07071 -0.0331 0.1178 1.0000
17.250 1.5637 0.08084 0.07527 -0.0342 0.1133 1.0000
17.500 1.5558 0.08529 0.07988 -0.0354 0.1081 1.0000
17.750 1.5432 0.09047 0.08513 -0.0370 0.1027 1.0000
18.000 1.5305 0.09584 0.09062 -0.0388 0.0967 1.0000
18.250 1.5163 0.10150 0.09636 -0.0408 0.0912 1.0000
18.500 1.5025 0.10722 0.10220 -0.0429 0.0854 1.0000
18.750 1.4839 0.11383 0.10886 -0.0456 0.0792 1.0000
19.000 1.4663 0.12047 0.11561 -0.0484 0.0714 1.0000
19.250 1.4481 0.12730 0.12251 -0.0515 0.0647 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 364 AIRFOIL (goe364-il)