GOE 385 AIRFOIL (goe385-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 385 AIRFOIL (goe385-il) Reynolds number: 500,000 Max Cl/Cd: 97.44 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe385-il-500000.txt Download as CSV file: xf-goe385-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 385 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3658 0.08732 0.08525 -0.0315 1.0000 0.0204
-8.000 -0.3632 0.08321 0.08116 -0.0357 1.0000 0.0204
-7.750 -0.3670 0.07836 0.07635 -0.0371 1.0000 0.0206
-7.500 -0.3681 0.07595 0.07397 -0.0356 0.9999 0.0208
-7.250 -0.3459 0.07218 0.07019 -0.0397 0.9957 0.0212
-7.000 -0.3208 0.06782 0.06579 -0.0459 0.9901 0.0217
-6.750 -0.2929 0.06290 0.06081 -0.0534 0.9835 0.0225
-6.500 -0.2661 0.05773 0.05554 -0.0603 0.9725 0.0239
-6.250 -0.2292 0.05050 0.04792 -0.0690 0.9581 0.0256
-6.000 -0.2170 0.04339 0.04054 -0.0712 0.9425 0.0261
-5.750 -0.2025 0.04116 0.03829 -0.0709 0.9257 0.0266
-5.500 -0.1850 0.03919 0.03623 -0.0706 0.9070 0.0272
-5.250 -0.1655 0.03697 0.03386 -0.0705 0.8880 0.0281
-5.000 -0.1436 0.03443 0.03106 -0.0704 0.8702 0.0298
-4.750 -0.1123 0.03300 0.02903 -0.0696 0.8539 0.0326
-4.500 -0.0953 0.02729 0.02303 -0.0702 0.8394 0.0340
-4.250 -0.0720 0.02590 0.02151 -0.0701 0.8245 0.0351
-4.000 -0.0471 0.02450 0.01992 -0.0700 0.8104 0.0367
-3.750 -0.0181 0.02424 0.01932 -0.0694 0.7954 0.0411
-3.500 0.0065 0.02094 0.01554 -0.0693 0.7804 0.0436
-3.250 0.0321 0.01973 0.01429 -0.0695 0.7622 0.0455
-3.000 0.0590 0.01882 0.01321 -0.0694 0.7430 0.0484
-2.750 0.0887 0.02011 0.01423 -0.0687 0.7242 0.0535
-2.500 0.1139 0.01667 0.01056 -0.0691 0.7092 0.0575
-2.250 0.1414 0.01599 0.00975 -0.0691 0.6947 0.0613
-2.000 0.1691 0.01528 0.00882 -0.0692 0.6817 0.0711
-1.750 0.1997 0.01360 0.00681 -0.0683 0.6711 0.0500
-1.500 0.2287 0.01259 0.00548 -0.0677 0.6611 0.0431
-1.250 0.2568 0.01212 0.00493 -0.0676 0.6515 0.0428
-1.000 0.2846 0.01152 0.00427 -0.0676 0.6430 0.0432
-0.750 0.3125 0.01112 0.00381 -0.0675 0.6348 0.0431
-0.500 0.3404 0.01070 0.00337 -0.0674 0.6272 0.0435
-0.250 0.3683 0.01039 0.00301 -0.0674 0.6196 0.0442
0.000 0.3965 0.01016 0.00275 -0.0675 0.6128 0.0452
0.250 0.4248 0.01000 0.00256 -0.0675 0.6054 0.0463
0.500 0.4531 0.00989 0.00244 -0.0676 0.5981 0.0485
0.750 0.4815 0.00980 0.00232 -0.0677 0.5895 0.0507
1.000 0.5099 0.00974 0.00221 -0.0678 0.5809 0.0529
1.250 0.5381 0.00969 0.00211 -0.0678 0.5721 0.0591
1.500 0.5535 0.00745 0.00226 -0.0657 0.5646 0.8964
1.750 0.5915 0.00741 0.00218 -0.0677 0.5556 1.0000
2.000 0.6196 0.00747 0.00219 -0.0677 0.5458 1.0000
2.250 0.6476 0.00755 0.00221 -0.0677 0.5357 1.0000
2.500 0.6754 0.00765 0.00224 -0.0677 0.5241 1.0000
2.750 0.7031 0.00775 0.00228 -0.0677 0.5107 1.0000
3.000 0.7308 0.00787 0.00233 -0.0677 0.4949 1.0000
3.250 0.7582 0.00801 0.00241 -0.0677 0.4763 1.0000
3.500 0.7853 0.00819 0.00250 -0.0677 0.4562 1.0000
3.750 0.8122 0.00841 0.00262 -0.0676 0.4353 1.0000
4.000 0.8388 0.00868 0.00277 -0.0675 0.4119 1.0000
4.250 0.8652 0.00896 0.00294 -0.0674 0.3897 1.0000
4.500 0.8916 0.00924 0.00313 -0.0673 0.3712 1.0000
4.750 0.9182 0.00950 0.00333 -0.0672 0.3575 1.0000
5.000 0.9447 0.00976 0.00355 -0.0671 0.3456 1.0000
5.250 0.9711 0.01004 0.00377 -0.0670 0.3343 1.0000
5.500 0.9977 0.01028 0.00400 -0.0669 0.3247 1.0000
5.750 1.0242 0.01052 0.00424 -0.0669 0.3160 1.0000
6.000 1.0501 0.01082 0.00449 -0.0667 0.3058 1.0000
6.250 1.0762 0.01107 0.00473 -0.0666 0.2938 1.0000
6.500 1.1023 0.01132 0.00497 -0.0665 0.2816 1.0000
6.750 1.1283 0.01158 0.00523 -0.0664 0.2713 1.0000
7.000 1.1538 0.01188 0.00551 -0.0662 0.2617 1.0000
7.250 1.1794 0.01216 0.00578 -0.0660 0.2499 1.0000
7.500 1.2048 0.01245 0.00608 -0.0658 0.2370 1.0000
7.750 1.2292 0.01283 0.00641 -0.0655 0.2208 1.0000
8.000 1.2532 0.01324 0.00677 -0.0652 0.1993 1.0000
8.250 1.2729 0.01409 0.00734 -0.0643 0.1515 1.0000
8.500 1.2897 0.01520 0.00821 -0.0631 0.1146 1.0000
8.750 1.3075 0.01617 0.00901 -0.0619 0.0779 1.0000
9.000 1.3198 0.01762 0.01020 -0.0600 0.0399 1.0000
9.250 1.3342 0.01879 0.01121 -0.0583 0.0220 1.0000
9.500 1.3517 0.01960 0.01209 -0.0570 0.0193 1.0000
9.750 1.3668 0.02056 0.01313 -0.0553 0.0174 1.0000
10.000 1.3797 0.02160 0.01427 -0.0534 0.0163 1.0000
10.250 1.3923 0.02250 0.01529 -0.0514 0.0157 1.0000
10.500 1.4004 0.02353 0.01642 -0.0488 0.0152 1.0000
10.750 1.4068 0.02474 0.01773 -0.0463 0.0147 1.0000
11.000 1.4117 0.02615 0.01924 -0.0439 0.0142 1.0000
11.250 1.4145 0.02783 0.02103 -0.0417 0.0138 1.0000
11.500 1.4145 0.02988 0.02319 -0.0397 0.0134 1.0000
11.750 1.4107 0.03250 0.02592 -0.0381 0.0130 1.0000
12.000 1.4020 0.03581 0.02936 -0.0367 0.0127 1.0000
12.250 1.3975 0.03890 0.03257 -0.0358 0.0124 1.0000
12.500 1.4016 0.04123 0.03501 -0.0355 0.0122 1.0000
12.750 1.4026 0.04398 0.03788 -0.0352 0.0120 1.0000
13.000 1.4016 0.04703 0.04104 -0.0351 0.0118 1.0000
13.250 1.4000 0.05020 0.04432 -0.0350 0.0116 1.0000
13.500 1.3975 0.05348 0.04771 -0.0350 0.0114 1.0000
13.750 1.3948 0.05682 0.05116 -0.0350 0.0112 1.0000
14.000 1.3919 0.06018 0.05462 -0.0351 0.0111 1.0000
14.250 1.3889 0.06359 0.05813 -0.0351 0.0109 1.0000
14.500 1.3864 0.06701 0.06164 -0.0353 0.0107 1.0000
14.750 1.3839 0.07049 0.06521 -0.0356 0.0106 1.0000
15.000 1.3814 0.07400 0.06882 -0.0359 0.0105 1.0000
15.250 1.3790 0.07756 0.07247 -0.0362 0.0103 1.0000
15.500 1.3768 0.08113 0.07613 -0.0367 0.0102 1.0000
15.750 1.3744 0.08473 0.07981 -0.0371 0.0100 1.0000
16.000 1.3721 0.08834 0.08352 -0.0376 0.0100 1.0000
16.250 1.3701 0.09185 0.08711 -0.0379 0.0098 1.0000
16.500 1.3684 0.09525 0.09059 -0.0380 0.0096 1.0000
16.750 1.3658 0.09862 0.09406 -0.0375 0.0095 1.0000
17.250 1.3466 0.10884 0.10461 -0.0402 0.0093 1.0000
17.500 1.3367 0.11467 0.11061 -0.0433 0.0092 1.0000
17.750 1.3268 0.12064 0.11674 -0.0466 0.0092 1.0000
18.000 1.3164 0.12686 0.12311 -0.0501 0.0091 1.0000
18.250 1.3062 0.13304 0.12945 -0.0534 0.0091 1.0000
18.500 1.2951 0.13963 0.13619 -0.0572 0.0091 1.0000
18.750 1.2838 0.14643 0.14314 -0.0613 0.0091 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 385 AIRFOIL (goe385-il)