GOE 562 AIRFOIL (goe562-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 562 AIRFOIL (goe562-il) Reynolds number: 1,000,000 Max Cl/Cd: 115.88 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe562-il-1000000-n5.txt Download as CSV file: xf-goe562-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 562 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3443 0.01816 0.01311 -0.1174 0.6102 0.0168
-9.000 -0.3291 0.01710 0.01182 -0.1153 0.6040 0.0173
-8.750 -0.3059 0.01705 0.01174 -0.1143 0.5968 0.0176
-8.500 -0.2805 0.01722 0.01191 -0.1137 0.5901 0.0178
-8.250 -0.2575 0.01714 0.01177 -0.1127 0.5839 0.0180
-8.000 -0.2332 0.01714 0.01175 -0.1119 0.5788 0.0183
-7.750 -0.2104 0.01699 0.01154 -0.1109 0.5731 0.0186
-7.250 -0.1639 0.01679 0.01123 -0.1090 0.5620 0.0194
-7.000 -0.1452 0.01604 0.01030 -0.1073 0.5564 0.0201
-6.500 -0.1035 0.01501 0.00899 -0.1045 0.5474 0.0210
-6.250 -0.0814 0.01465 0.00851 -0.1033 0.5427 0.0212
-6.000 -0.0621 0.01395 0.00767 -0.1017 0.5371 0.0217
-5.750 -0.0386 0.01384 0.00754 -0.1007 0.5329 0.0221
-5.500 -0.0154 0.01365 0.00729 -0.0997 0.5275 0.0223
-5.250 0.0077 0.01348 0.00707 -0.0987 0.5228 0.0226
-5.000 0.0308 0.01328 0.00680 -0.0977 0.5185 0.0229
-4.750 0.0540 0.01299 0.00645 -0.0967 0.5144 0.0233
-4.500 0.0773 0.01276 0.00615 -0.0957 0.5096 0.0238
-4.250 0.0999 0.01242 0.00572 -0.0945 0.5052 0.0241
-4.000 0.1232 0.01213 0.00535 -0.0935 0.5001 0.0245
-3.750 0.1458 0.01186 0.00499 -0.0924 0.4919 0.0248
-3.500 0.1685 0.01160 0.00466 -0.0912 0.4876 0.0252
-3.250 0.1917 0.01138 0.00439 -0.0902 0.4834 0.0255
-3.000 0.2148 0.01121 0.00417 -0.0891 0.4790 0.0258
-2.750 0.2377 0.01112 0.00403 -0.0880 0.4739 0.0261
-2.500 0.2604 0.01095 0.00381 -0.0869 0.4697 0.0262
-2.250 0.2807 0.01056 0.00337 -0.0852 0.4649 0.0268
-2.000 0.3020 0.01037 0.00313 -0.0838 0.4598 0.0270
-1.750 0.3241 0.01025 0.00297 -0.0825 0.4550 0.0275
-1.500 0.3469 0.01015 0.00286 -0.0814 0.4504 0.0281
-1.250 0.3693 0.01004 0.00272 -0.0802 0.4467 0.0284
-1.000 0.3916 0.00997 0.00262 -0.0789 0.4428 0.0287
-0.750 0.4145 0.00990 0.00253 -0.0779 0.4395 0.0292
-0.500 0.4374 0.00983 0.00244 -0.0768 0.4351 0.0298
-0.250 0.4601 0.00978 0.00236 -0.0757 0.4306 0.0304
0.000 0.4825 0.00978 0.00232 -0.0745 0.4256 0.0309
0.250 0.5059 0.00974 0.00227 -0.0735 0.4221 0.0316
0.500 0.5287 0.00974 0.00224 -0.0725 0.4160 0.0321
0.750 0.5511 0.00975 0.00222 -0.0713 0.4115 0.0325
1.000 0.5737 0.00971 0.00216 -0.0702 0.4059 0.0339
1.250 0.5963 0.00972 0.00216 -0.0691 0.4010 0.0355
1.500 0.6186 0.00976 0.00218 -0.0680 0.3969 0.0363
1.750 0.6414 0.00979 0.00220 -0.0670 0.3938 0.0383
2.000 0.6644 0.00981 0.00223 -0.0660 0.3901 0.0402
2.250 0.6871 0.00984 0.00227 -0.0650 0.3862 0.0455
2.500 0.7092 0.00990 0.00232 -0.0638 0.3821 0.0520
2.750 0.7305 0.00982 0.00242 -0.0626 0.3785 0.1364
3.000 0.7506 0.00953 0.00257 -0.0613 0.3744 0.3377
3.500 1.0939 0.00953 0.00386 -0.1278 0.3477 1.0000
3.750 1.1149 0.00966 0.00397 -0.1265 0.3412 1.0000
4.000 1.1351 0.00982 0.00409 -0.1250 0.3349 1.0000
4.250 1.1553 0.00997 0.00421 -0.1236 0.3270 1.0000
4.500 1.1744 0.01016 0.00436 -0.1220 0.3185 1.0000
4.750 1.1928 0.01037 0.00452 -0.1202 0.3084 1.0000
5.000 1.2069 0.01074 0.00476 -0.1176 0.2876 1.0000
5.250 1.2207 0.01104 0.00499 -0.1149 0.2741 1.0000
5.500 1.2347 0.01128 0.00519 -0.1122 0.2648 1.0000
5.750 1.2499 0.01148 0.00538 -0.1098 0.2587 1.0000
6.000 1.2647 0.01171 0.00558 -0.1073 0.2517 1.0000
6.250 1.2780 0.01200 0.00583 -0.1046 0.2426 1.0000
6.500 1.2921 0.01227 0.00606 -0.1020 0.2359 1.0000
6.750 1.3060 0.01254 0.00632 -0.0994 0.2298 1.0000
7.000 1.3216 0.01276 0.00654 -0.0973 0.2257 1.0000
7.250 1.3361 0.01305 0.00681 -0.0949 0.2207 1.0000
7.500 1.3494 0.01337 0.00712 -0.0923 0.2145 1.0000
7.750 1.3637 0.01367 0.00741 -0.0900 0.2089 1.0000
8.000 1.3776 0.01400 0.00773 -0.0876 0.2043 1.0000
8.250 1.3912 0.01434 0.00808 -0.0852 0.1991 1.0000
8.500 1.4043 0.01471 0.00843 -0.0828 0.1921 1.0000
8.750 1.4154 0.01516 0.00886 -0.0801 0.1844 1.0000
9.000 1.4087 0.01631 0.00980 -0.0745 0.1486 1.0000
9.250 1.3808 0.01846 0.01172 -0.0659 0.0976 1.0000
9.500 1.3630 0.02048 0.01361 -0.0598 0.0598 1.0000
9.750 1.3668 0.02161 0.01473 -0.0570 0.0499 1.0000
10.000 1.3707 0.02282 0.01596 -0.0545 0.0413 1.0000
10.250 1.3635 0.02487 0.01793 -0.0511 0.0203 1.0000
10.500 1.3662 0.02643 0.01952 -0.0491 0.0157 1.0000
10.750 1.3722 0.02784 0.02098 -0.0475 0.0139 1.0000
11.000 1.3765 0.02945 0.02264 -0.0460 0.0124 1.0000
11.250 1.3807 0.03117 0.02441 -0.0446 0.0114 1.0000
11.500 1.3860 0.03286 0.02616 -0.0435 0.0111 1.0000
11.750 1.3903 0.03469 0.02806 -0.0424 0.0105 1.0000
12.000 1.3934 0.03668 0.03011 -0.0414 0.0099 1.0000
12.250 1.3950 0.03889 0.03238 -0.0405 0.0093 1.0000
12.500 1.3944 0.04134 0.03489 -0.0396 0.0086 1.0000
12.750 1.3937 0.04387 0.03749 -0.0387 0.0085 1.0000
13.000 1.3942 0.04629 0.03998 -0.0380 0.0082 1.0000
13.250 1.3936 0.04887 0.04264 -0.0373 0.0081 1.0000
13.500 1.3921 0.05160 0.04543 -0.0367 0.0077 1.0000
13.750 1.3908 0.05434 0.04824 -0.0362 0.0073 1.0000
14.000 1.3881 0.05728 0.05126 -0.0357 0.0073 1.0000
14.250 1.3847 0.06037 0.05440 -0.0353 0.0069 1.0000
14.500 1.3810 0.06354 0.05765 -0.0350 0.0066 1.0000
14.750 1.3764 0.06687 0.06105 -0.0348 0.0065 1.0000
15.000 1.3717 0.07025 0.06451 -0.0346 0.0065 1.0000
15.250 1.3644 0.07403 0.06837 -0.0345 0.0061 1.0000
15.500 1.3594 0.07758 0.07199 -0.0345 0.0061 1.0000
15.750 1.3551 0.08104 0.07553 -0.0346 0.0059 1.0000
16.000 1.3499 0.08466 0.07923 -0.0347 0.0059 1.0000
16.250 1.3414 0.08881 0.08347 -0.0349 0.0058 1.0000
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