GOE 562 AIRFOIL (goe562-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 562 AIRFOIL (goe562-il) Reynolds number: 500,000 Max Cl/Cd: 101.4 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe562-il-500000.txt Download as CSV file: xf-goe562-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 0.0263 0.08689 0.08458 -0.0919 0.8884 0.0236
-10.250 0.0620 0.07930 0.07680 -0.1055 0.8670 0.0258
-10.000 -0.1473 0.10123 0.09911 -0.0616 0.9211 0.0222
-9.750 -0.1010 0.09516 0.09294 -0.0732 0.9069 0.0227
-9.500 -0.0362 0.08731 0.08492 -0.0918 0.8899 0.0246
-9.250 0.0094 0.07734 0.07468 -0.1166 0.8552 0.0262
-9.000 0.0435 0.07206 0.06909 -0.1246 0.8048 0.0266
-8.750 0.0601 0.07001 0.06680 -0.1259 0.7621 0.0269
-8.500 0.0688 0.06789 0.06452 -0.1268 0.7314 0.0274
-8.250 0.0743 0.06556 0.06207 -0.1279 0.7088 0.0280
-8.000 0.0750 0.06284 0.05927 -0.1296 0.6917 0.0289
-7.750 0.0702 0.06003 0.05640 -0.1305 0.6781 0.0294
-7.500 0.0622 0.05694 0.05324 -0.1309 0.6676 0.0304
-7.250 0.0489 0.05090 0.04698 -0.1335 0.6601 0.0316
-7.000 0.0379 0.04600 0.04191 -0.1315 0.6530 0.0320
-6.750 0.0456 0.04419 0.04005 -0.1300 0.6441 0.0323
-6.500 0.0554 0.04282 0.03863 -0.1285 0.6360 0.0327
-6.250 0.0654 0.04132 0.03705 -0.1269 0.6286 0.0333
-6.000 0.0751 0.03952 0.03513 -0.1252 0.6222 0.0340
-5.750 0.0843 0.03745 0.03294 -0.1232 0.6159 0.0354
-5.500 0.0911 0.03542 0.03031 -0.1188 0.6109 0.0384
-5.250 0.0893 0.03088 0.02562 -0.1155 0.6063 0.0393
-5.000 0.1047 0.02960 0.02434 -0.1140 0.6003 0.0400
-4.750 0.1206 0.02855 0.02319 -0.1124 0.5947 0.0409
-4.500 0.1358 0.02735 0.02186 -0.1103 0.5893 0.0423
-4.250 0.1559 0.02750 0.02164 -0.1075 0.5835 0.0466
-4.000 0.1584 0.02434 0.01808 -0.1030 0.5792 0.0478
-3.750 0.1767 0.02276 0.01649 -0.1018 0.5744 0.0489
-3.500 0.1953 0.02184 0.01554 -0.1002 0.5691 0.0503
-3.250 0.2147 0.02111 0.01469 -0.0986 0.5641 0.0528
-3.000 0.2406 0.02224 0.01554 -0.0972 0.5592 0.0571
-2.750 0.2503 0.01919 0.01228 -0.0941 0.5547 0.0597
-2.500 0.2710 0.01832 0.01138 -0.0928 0.5493 0.0615
-2.250 0.2928 0.01779 0.01071 -0.0915 0.5441 0.0643
-2.000 0.3184 0.01561 0.00800 -0.0892 0.5399 0.0454
-1.750 0.3420 0.01478 0.00706 -0.0881 0.5344 0.0448
-1.500 0.3667 0.01425 0.00640 -0.0873 0.5295 0.0448
-1.250 0.3910 0.01375 0.00584 -0.0864 0.5246 0.0449
-1.000 0.4144 0.01349 0.00553 -0.0854 0.5192 0.0458
-0.750 0.4385 0.01326 0.00521 -0.0845 0.5145 0.0465
-0.500 0.4623 0.01297 0.00490 -0.0836 0.5101 0.0467
-0.250 0.4849 0.01259 0.00451 -0.0824 0.5057 0.0473
0.000 0.5055 0.01217 0.00406 -0.0808 0.5013 0.0486
0.250 0.5273 0.01201 0.00386 -0.0795 0.4967 0.0500
0.500 0.5488 0.01185 0.00372 -0.0782 0.4923 0.0520
0.750 0.5705 0.01175 0.00361 -0.0768 0.4876 0.0545
1.000 0.5924 0.01171 0.00351 -0.0756 0.4831 0.0561
1.250 0.6142 0.01161 0.00340 -0.0742 0.4787 0.0590
1.500 0.6355 0.01150 0.00330 -0.0728 0.4745 0.0652
1.750 0.6568 0.01144 0.00324 -0.0714 0.4705 0.0803
2.000 0.9338 0.01048 0.00432 -0.1282 0.4581 1.0000
2.250 0.9553 0.01059 0.00437 -0.1270 0.4539 1.0000
2.500 0.9770 0.01072 0.00445 -0.1258 0.4500 1.0000
2.750 0.9992 0.01078 0.00453 -0.1247 0.4460 1.0000
3.000 1.0207 0.01088 0.00460 -0.1235 0.4417 1.0000
3.250 1.0419 0.01102 0.00469 -0.1222 0.4378 1.0000
3.500 1.0633 0.01115 0.00479 -0.1210 0.4339 1.0000
3.750 1.0850 0.01123 0.00490 -0.1199 0.4298 1.0000
4.000 1.1059 0.01135 0.00500 -0.1185 0.4255 1.0000
4.250 1.1260 0.01152 0.00512 -0.1171 0.4212 1.0000
4.500 1.1471 0.01163 0.00525 -0.1158 0.4171 1.0000
4.750 1.1677 0.01173 0.00538 -0.1145 0.4124 1.0000
5.000 1.1871 0.01189 0.00552 -0.1129 0.4079 1.0000
5.250 1.2065 0.01205 0.00567 -0.1113 0.4034 1.0000
5.500 1.2265 0.01216 0.00582 -0.1098 0.3983 1.0000
5.750 1.2447 0.01232 0.00597 -0.1080 0.3931 1.0000
6.000 1.2617 0.01249 0.00614 -0.1060 0.3880 1.0000
6.250 1.2786 0.01261 0.00630 -0.1038 0.3821 1.0000
6.500 1.2920 0.01281 0.00647 -0.1011 0.3756 1.0000
6.750 1.3078 0.01294 0.00663 -0.0988 0.3670 1.0000
7.000 1.3208 0.01317 0.00683 -0.0960 0.3582 1.0000
7.250 1.3333 0.01342 0.00705 -0.0931 0.3463 1.0000
7.500 1.3468 0.01368 0.00730 -0.0905 0.3345 1.0000
7.750 1.3591 0.01400 0.00759 -0.0877 0.3224 1.0000
8.000 1.3703 0.01438 0.00793 -0.0848 0.3108 1.0000
8.250 1.3804 0.01481 0.00833 -0.0817 0.2995 1.0000
8.500 1.3902 0.01529 0.00876 -0.0787 0.2878 1.0000
8.750 1.4016 0.01573 0.00921 -0.0760 0.2791 1.0000
9.000 1.4107 0.01628 0.00973 -0.0730 0.2714 1.0000
9.250 1.4206 0.01681 0.01026 -0.0702 0.2623 1.0000
9.500 1.4292 0.01743 0.01086 -0.0673 0.2544 1.0000
9.750 1.4393 0.01801 0.01146 -0.0647 0.2483 1.0000
10.000 1.4467 0.01873 0.01218 -0.0619 0.2414 1.0000
10.250 1.4553 0.01944 0.01291 -0.0594 0.2339 1.0000
10.500 1.4617 0.02031 0.01379 -0.0566 0.2272 1.0000
10.750 1.4650 0.02140 0.01486 -0.0538 0.2166 1.0000
11.000 1.4698 0.02250 0.01597 -0.0513 0.2053 1.0000
11.250 1.4685 0.02406 0.01749 -0.0484 0.1893 1.0000
11.500 1.4500 0.02697 0.02019 -0.0444 0.1450 1.0000
11.750 1.4200 0.03117 0.02422 -0.0404 0.1071 1.0000
12.000 1.3911 0.03582 0.02880 -0.0375 0.0742 1.0000
12.250 1.3741 0.03970 0.03265 -0.0357 0.0572 1.0000
12.500 1.3611 0.04337 0.03635 -0.0344 0.0440 1.0000
12.750 1.3458 0.04739 0.04036 -0.0332 0.0294 1.0000
13.000 1.3356 0.05100 0.04398 -0.0324 0.0240 1.0000
13.250 1.3271 0.05454 0.04759 -0.0317 0.0216 1.0000
13.500 1.3201 0.05801 0.05114 -0.0312 0.0201 1.0000
13.750 1.3146 0.06138 0.05461 -0.0308 0.0195 1.0000
14.000 1.3086 0.06489 0.05822 -0.0305 0.0189 1.0000
14.250 1.3032 0.06838 0.06181 -0.0303 0.0186 1.0000
14.500 1.2932 0.07252 0.06605 -0.0303 0.0178 1.0000
14.750 1.2836 0.07671 0.07035 -0.0303 0.0175 1.0000
15.000 1.2734 0.08108 0.07481 -0.0305 0.0170 1.0000
15.250 1.2624 0.08559 0.07942 -0.0308 0.0168 1.0000
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