GOE 654 AIRFOIL (goe654-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 654 AIRFOIL (goe654-il) Reynolds number: 1,000,000 Max Cl/Cd: 115.31 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe654-il-1000000-n5.txt Download as CSV file: xf-goe654-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 654 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
------ -------- --------- --------- -------- -------- --------
-11.750 -0.4318 0.03810 0.03554 -0.1242 0.8784 0.0191
-11.500 -0.4849 0.03256 0.02951 -0.1244 0.8269 0.0193
-11.250 -0.5104 0.02951 0.02611 -0.1206 0.8027 0.0194
-11.000 -0.5135 0.02734 0.02364 -0.1183 0.7852 0.0196
-10.750 -0.5083 0.02564 0.02166 -0.1163 0.7705 0.0197
-10.500 -0.4970 0.02435 0.02016 -0.1147 0.7569 0.0197
-10.250 -0.4828 0.02323 0.01884 -0.1133 0.7438 0.0198
-10.000 -0.4720 0.02149 0.01688 -0.1117 0.7319 0.0200
-9.750 -0.4542 0.02056 0.01579 -0.1106 0.7203 0.0202
-9.500 -0.4340 0.01991 0.01503 -0.1096 0.7079 0.0203
-9.250 -0.4128 0.01930 0.01431 -0.1087 0.6957 0.0204
-9.000 -0.3913 0.01871 0.01359 -0.1078 0.6833 0.0205
-8.750 -0.3691 0.01822 0.01299 -0.1070 0.6696 0.0206
-8.500 -0.3466 0.01774 0.01238 -0.1061 0.6544 0.0208
-8.250 -0.3239 0.01727 0.01178 -0.1053 0.6375 0.0210
-8.000 -0.3002 0.01696 0.01135 -0.1046 0.6213 0.0212
-7.750 -0.2771 0.01646 0.01071 -0.1038 0.6060 0.0214
-7.250 -0.2301 0.01548 0.00946 -0.1023 0.5762 0.0217
-7.000 -0.2061 0.01502 0.00887 -0.1016 0.5635 0.0219
-6.750 -0.1817 0.01459 0.00833 -0.1009 0.5531 0.0220
-6.500 -0.1569 0.01418 0.00782 -0.1003 0.5444 0.0222
-6.250 -0.1318 0.01379 0.00733 -0.0997 0.5374 0.0224
-6.000 -0.1062 0.01341 0.00688 -0.0992 0.5319 0.0225
-5.750 -0.0808 0.01307 0.00646 -0.0987 0.5253 0.0227
-5.500 -0.0550 0.01277 0.00608 -0.0982 0.5197 0.0229
-5.250 -0.0290 0.01245 0.00571 -0.0977 0.5148 0.0230
-5.000 -0.0030 0.01218 0.00538 -0.0973 0.5095 0.0232
-4.750 0.0230 0.01195 0.00508 -0.0968 0.5044 0.0233
-4.500 0.0497 0.01171 0.00481 -0.0964 0.5013 0.0234
-4.250 0.0767 0.01154 0.00461 -0.0961 0.4977 0.0236
-4.000 0.1020 0.01111 0.00414 -0.0955 0.4943 0.0240
-3.750 0.1281 0.01086 0.00387 -0.0951 0.4908 0.0243
-3.500 0.1545 0.01068 0.00365 -0.0947 0.4872 0.0246
-3.250 0.1816 0.01050 0.00346 -0.0944 0.4841 0.0249
-3.000 0.2086 0.01034 0.00329 -0.0941 0.4802 0.0251
-2.750 0.2355 0.01021 0.00313 -0.0938 0.4757 0.0254
-2.500 0.2623 0.01009 0.00298 -0.0934 0.4718 0.0257
-2.250 0.2894 0.00997 0.00284 -0.0931 0.4685 0.0261
-2.000 0.3167 0.00986 0.00271 -0.0929 0.4648 0.0264
-1.750 0.3439 0.00976 0.00260 -0.0926 0.4603 0.0267
-1.500 0.3708 0.00969 0.00249 -0.0923 0.4552 0.0271
-1.250 0.3981 0.00962 0.00240 -0.0920 0.4505 0.0274
-1.000 0.4253 0.00956 0.00232 -0.0918 0.4439 0.0277
-0.750 0.4521 0.00954 0.00226 -0.0914 0.4374 0.0280
-0.500 0.4792 0.00945 0.00216 -0.0912 0.4311 0.0289
-0.250 0.5058 0.00944 0.00211 -0.0908 0.4229 0.0299
0.000 0.5327 0.00942 0.00208 -0.0905 0.4142 0.0308
0.250 0.5592 0.00946 0.00206 -0.0902 0.4036 0.0316
0.500 0.5852 0.00952 0.00206 -0.0897 0.3906 0.0326
0.750 0.6110 0.00960 0.00208 -0.0893 0.3766 0.0341
1.000 0.6368 0.00968 0.00211 -0.0888 0.3638 0.0373
1.250 0.6617 0.00961 0.00215 -0.0882 0.3538 0.0950
1.500 0.6859 0.00948 0.00222 -0.0876 0.3451 0.1952
1.750 0.7091 0.00916 0.00231 -0.0869 0.3390 0.3608
2.000 0.7335 0.00907 0.00241 -0.0863 0.3326 0.4422
2.250 0.7575 0.00901 0.00253 -0.0856 0.3269 0.5235
2.500 0.7812 0.00889 0.00264 -0.0848 0.3221 0.6141
2.750 0.8026 0.00871 0.00276 -0.0835 0.3175 0.7290
3.250 0.9097 0.00861 0.00314 -0.0947 0.3076 1.0000
3.500 0.9336 0.00874 0.00323 -0.0938 0.3042 1.0000
3.750 0.9571 0.00889 0.00334 -0.0930 0.3001 1.0000
4.000 0.9804 0.00904 0.00346 -0.0921 0.2965 1.0000
4.250 1.0045 0.00917 0.00357 -0.0913 0.2939 1.0000
4.500 1.0286 0.00928 0.00368 -0.0906 0.2913 1.0000
4.750 1.0525 0.00942 0.00380 -0.0898 0.2879 1.0000
5.000 1.0757 0.00959 0.00394 -0.0889 0.2837 1.0000
5.250 1.0985 0.00978 0.00409 -0.0880 0.2789 1.0000
5.500 1.1226 0.00991 0.00422 -0.0873 0.2755 1.0000
5.750 1.1460 0.01007 0.00436 -0.0865 0.2713 1.0000
6.000 1.1690 0.01024 0.00452 -0.0856 0.2676 1.0000
6.250 1.1917 0.01043 0.00469 -0.0847 0.2644 1.0000
6.500 1.2151 0.01058 0.00485 -0.0840 0.2624 1.0000
6.750 1.2384 0.01074 0.00500 -0.0832 0.2588 1.0000
7.000 1.2603 0.01095 0.00519 -0.0822 0.2529 1.0000
7.250 1.2815 0.01119 0.00540 -0.0811 0.2482 1.0000
7.500 1.3039 0.01136 0.00558 -0.0802 0.2435 1.0000
7.750 1.3246 0.01160 0.00580 -0.0790 0.2375 1.0000
8.000 1.3439 0.01186 0.00603 -0.0776 0.2317 1.0000
8.250 1.3614 0.01213 0.00628 -0.0758 0.2234 1.0000
8.500 1.3777 0.01244 0.00655 -0.0739 0.2159 1.0000
8.750 1.3934 0.01279 0.00686 -0.0719 0.2071 1.0000
9.000 1.4087 0.01317 0.00720 -0.0699 0.1979 1.0000
9.250 1.4212 0.01368 0.00765 -0.0675 0.1853 1.0000
9.500 1.4306 0.01433 0.00822 -0.0647 0.1698 1.0000
9.750 1.4400 0.01503 0.00884 -0.0621 0.1562 1.0000
10.000 1.4482 0.01580 0.00954 -0.0594 0.1430 1.0000
10.250 1.4558 0.01666 0.01034 -0.0569 0.1304 1.0000
10.500 1.4628 0.01760 0.01123 -0.0544 0.1186 1.0000
10.750 1.4637 0.01896 0.01250 -0.0515 0.1025 1.0000
11.000 1.4549 0.02106 0.01444 -0.0480 0.0758 1.0000
11.250 1.4320 0.02447 0.01769 -0.0442 0.0396 1.0000
11.500 1.4213 0.02738 0.02057 -0.0420 0.0173 1.0000
11.750 1.4283 0.02900 0.02223 -0.0411 0.0154 1.0000
12.000 1.4357 0.03066 0.02393 -0.0404 0.0141 1.0000
12.250 1.4432 0.03233 0.02566 -0.0398 0.0134 1.0000
12.500 1.4493 0.03419 0.02757 -0.0393 0.0128 1.0000
12.750 1.4536 0.03627 0.02971 -0.0387 0.0121 1.0000
13.000 1.4570 0.03848 0.03198 -0.0383 0.0117 1.0000
13.250 1.4610 0.04067 0.03424 -0.0380 0.0113 1.0000
13.500 1.4645 0.04299 0.03663 -0.0378 0.0110 1.0000
13.750 1.4665 0.04551 0.03921 -0.0376 0.0105 1.0000
14.000 1.4675 0.04818 0.04195 -0.0375 0.0102 1.0000
14.250 1.4672 0.05107 0.04490 -0.0375 0.0100 1.0000
14.500 1.4656 0.05413 0.04803 -0.0376 0.0097 1.0000
14.750 1.4628 0.05736 0.05134 -0.0377 0.0094 1.0000
15.000 1.4592 0.06071 0.05476 -0.0378 0.0092 1.0000
15.250 1.4549 0.06418 0.05831 -0.0380 0.0090 1.0000
15.500 1.4515 0.06759 0.06180 -0.0383 0.0089 1.0000
15.750 1.4469 0.07116 0.06544 -0.0386 0.0087 1.0000
16.000 1.4417 0.07486 0.06923 -0.0390 0.0087 1.0000
16.250 1.4365 0.07860 0.07304 -0.0395 0.0085 1.0000
16.500 1.4309 0.08244 0.07696 -0.0400 0.0083 1.0000
16.750 1.4233 0.08657 0.08118 -0.0407 0.0082 1.0000
17.000 1.4152 0.09084 0.08554 -0.0414 0.0081 1.0000
17.250 1.4087 0.09497 0.08974 -0.0422 0.0079 1.0000
17.500 1.4007 0.09933 0.09418 -0.0431 0.0078 1.0000
17.750 1.3919 0.10383 0.09877 -0.0441 0.0077 1.0000
18.000 1.3850 0.10812 0.10313 -0.0452 0.0075 1.0000
18.250 1.3754 0.11292 0.10801 -0.0465 0.0075 1.0000
18.500 1.3670 0.11757 0.11274 -0.0479 0.0073 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 654 AIRFOIL (goe654-il)