GOE 602 AIRFOIL (goe602-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 602 AIRFOIL (goe602-il) Reynolds number: 200,000 Max Cl/Cd: 77.49 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe602-il-200000.txt Download as CSV file: xf-goe602-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 602 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 -0.3943 0.09740 0.09402 -0.0414 1.0000 0.0457
-8.750 -0.4058 0.09482 0.09151 -0.0415 1.0000 0.0458
-8.500 -0.4200 0.09254 0.08932 -0.0405 1.0000 0.0459
-8.250 -0.4258 0.08750 0.08434 -0.0374 1.0000 0.0469
-8.000 -0.4253 0.08558 0.08244 -0.0341 1.0000 0.0478
-7.750 -0.4366 0.08396 0.08088 -0.0312 1.0000 0.0483
-7.500 -0.4496 0.08200 0.07898 -0.0292 1.0000 0.0487
-7.250 -0.4577 0.07964 0.07666 -0.0282 1.0000 0.0495
-7.000 -0.4652 0.07692 0.07398 -0.0280 1.0000 0.0503
-6.750 -0.4713 0.07380 0.07087 -0.0284 1.0000 0.0514
-6.500 -0.4735 0.07006 0.06712 -0.0302 0.9997 0.0530
-6.250 -0.4417 0.05713 0.05355 -0.0500 0.9915 0.0590
-6.000 -0.4187 0.05438 0.05096 -0.0507 0.9879 0.0610
-5.750 -0.3884 0.05114 0.04761 -0.0545 0.9841 0.0656
-5.500 -0.3634 0.04555 0.04163 -0.0595 0.9772 0.0743
-5.250 -0.3298 0.04179 0.03736 -0.0641 0.9724 0.0863
-5.000 -0.3028 0.03901 0.03467 -0.0655 0.9677 0.0899
-4.750 -0.2714 0.02742 0.02145 -0.0663 0.9630 0.0504
-4.500 -0.2350 0.02499 0.01857 -0.0682 0.9595 0.0523
-4.250 -0.1955 0.02243 0.01549 -0.0702 0.9572 0.0523
-4.000 -0.1681 0.02119 0.01397 -0.0700 0.9496 0.0549
-3.750 -0.1286 0.02008 0.01256 -0.0720 0.9459 0.0575
-3.500 -0.0881 0.01842 0.01076 -0.0744 0.9436 0.0601
-3.250 -0.0606 0.01766 0.01001 -0.0743 0.9355 0.0646
-3.000 -0.0209 0.01683 0.00912 -0.0765 0.9318 0.0684
-2.750 0.0212 0.01593 0.00821 -0.0792 0.9293 0.0736
-2.500 0.0478 0.01544 0.00776 -0.0789 0.9205 0.0812
-2.250 0.0887 0.01471 0.00704 -0.0814 0.9172 0.0963
-2.000 0.1277 0.01306 0.00644 -0.0842 0.9150 0.3364
-1.750 0.1493 0.01230 0.00633 -0.0829 0.9056 0.5065
-1.500 0.1825 0.01145 0.00614 -0.0834 0.9021 0.7032
-1.250 0.2358 0.01082 0.00603 -0.0871 0.9012 0.9200
-1.000 0.3242 0.01055 0.00563 -0.0995 0.9023 1.0000
-0.750 0.3521 0.01041 0.00538 -0.0994 0.8925 1.0000
-0.500 0.3908 0.01015 0.00500 -0.1013 0.8865 1.0000
-0.250 0.4157 0.01006 0.00481 -0.1005 0.8749 1.0000
0.000 0.4431 0.00995 0.00461 -0.1001 0.8640 1.0000
0.250 0.4726 0.00982 0.00440 -0.1001 0.8537 1.0000
0.500 0.5021 0.00969 0.00418 -0.1001 0.8429 1.0000
0.750 0.5270 0.00963 0.00406 -0.0992 0.8293 1.0000
1.000 0.5523 0.00959 0.00394 -0.0984 0.8153 1.0000
1.250 0.5780 0.00956 0.00384 -0.0976 0.8006 1.0000
1.500 0.6041 0.00955 0.00375 -0.0969 0.7853 1.0000
1.750 0.6278 0.00959 0.00372 -0.0958 0.7673 1.0000
2.000 0.6523 0.00964 0.00368 -0.0948 0.7482 1.0000
2.250 0.6775 0.00971 0.00365 -0.0940 0.7287 1.0000
2.500 0.7010 0.00984 0.00369 -0.0929 0.7084 1.0000
2.750 0.7248 0.00998 0.00375 -0.0918 0.6884 1.0000
3.000 0.7490 0.01016 0.00383 -0.0909 0.6701 1.0000
3.250 0.7721 0.01035 0.00398 -0.0898 0.6506 1.0000
3.500 0.7953 0.01055 0.00412 -0.0888 0.6311 1.0000
3.750 0.8186 0.01078 0.00426 -0.0877 0.6125 1.0000
4.000 0.8416 0.01100 0.00447 -0.0867 0.5940 1.0000
4.250 0.8646 0.01124 0.00468 -0.0856 0.5760 1.0000
4.500 0.8876 0.01150 0.00490 -0.0846 0.5582 1.0000
4.750 0.9101 0.01177 0.00512 -0.0835 0.5398 1.0000
5.000 0.9314 0.01202 0.00537 -0.0821 0.5173 1.0000
5.250 0.9509 0.01231 0.00557 -0.0804 0.4895 1.0000
5.500 0.9697 0.01263 0.00579 -0.0786 0.4594 1.0000
5.750 0.9883 0.01298 0.00606 -0.0769 0.4288 1.0000
6.000 1.0077 0.01335 0.00637 -0.0753 0.4012 1.0000
6.250 1.0259 0.01376 0.00672 -0.0735 0.3691 1.0000
6.500 1.0420 0.01428 0.00709 -0.0714 0.3259 1.0000
6.750 1.0542 0.01507 0.00756 -0.0688 0.2654 1.0000
7.000 1.0639 0.01617 0.00822 -0.0659 0.1914 1.0000
7.250 1.0667 0.01793 0.00929 -0.0622 0.0942 1.0000
7.500 1.0746 0.01934 0.01034 -0.0591 0.0454 1.0000
7.750 1.0873 0.02030 0.01133 -0.0565 0.0386 1.0000
8.000 1.0982 0.02125 0.01237 -0.0537 0.0357 1.0000
8.250 1.1051 0.02240 0.01365 -0.0503 0.0337 1.0000
8.500 1.1131 0.02347 0.01484 -0.0472 0.0323 1.0000
8.750 1.1210 0.02459 0.01607 -0.0443 0.0308 1.0000
9.000 1.1276 0.02586 0.01744 -0.0413 0.0296 1.0000
9.250 1.1340 0.02725 0.01893 -0.0385 0.0288 1.0000
9.500 1.1412 0.02876 0.02051 -0.0359 0.0281 1.0000
9.750 1.1504 0.03036 0.02217 -0.0337 0.0275 1.0000
10.000 1.1633 0.03209 0.02396 -0.0319 0.0270 1.0000
10.250 1.1825 0.03410 0.02599 -0.0309 0.0263 1.0000
10.500 1.2333 0.03912 0.03111 -0.0346 0.0247 1.0000
10.750 1.2608 0.04221 0.03447 -0.0351 0.0248 1.0000
11.000 1.2756 0.04419 0.03665 -0.0335 0.0250 1.0000
11.250 1.2848 0.04590 0.03858 -0.0312 0.0253 1.0000
11.500 1.2882 0.04737 0.04029 -0.0280 0.0257 1.0000
11.750 1.2845 0.04902 0.04236 -0.0241 0.0271 1.0000
12.000 1.2701 0.05325 0.04721 -0.0199 0.0298 1.0000
12.250 1.2698 0.06285 0.05757 -0.0176 0.0415 1.0000
12.750 1.1583 0.06038 0.05556 -0.0091 0.0357 1.0000
13.000 1.1191 0.06671 0.06227 -0.0083 0.0368 1.0000
13.250 1.0855 0.07344 0.06930 -0.0090 0.0376 1.0000
13.500 1.0558 0.07985 0.07593 -0.0106 0.0378 1.0000
13.750 1.0275 0.08616 0.08242 -0.0130 0.0376 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 602 AIRFOIL (goe602-il)