GOE 527 AIRFOIL (goe527-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 527 AIRFOIL (goe527-il) Reynolds number: 500,000 Max Cl/Cd: 96.2 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe527-il-500000-n5.txt Download as CSV file: xf-goe527-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 527 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-14.500 -0.7041 0.04807 0.04454 -0.1064 0.9843 0.0267
-14.250 -0.7136 0.03980 0.03603 -0.1174 0.9703 0.0268
-14.000 -0.7036 0.03440 0.03040 -0.1269 0.9558 0.0270
-13.750 -0.6772 0.03109 0.02691 -0.1347 0.9430 0.0273
-13.500 -0.6458 0.02873 0.02435 -0.1408 0.9278 0.0276
-13.250 -0.6209 0.02689 0.02228 -0.1442 0.9083 0.0279
-13.000 -0.6081 0.02560 0.02078 -0.1439 0.8864 0.0282
-12.750 -0.6023 0.02472 0.01971 -0.1415 0.8661 0.0285
-12.250 -0.5830 0.02313 0.01776 -0.1373 0.8350 0.0290
-12.000 -0.5700 0.02242 0.01687 -0.1354 0.8217 0.0293
-11.750 -0.5556 0.02177 0.01603 -0.1336 0.8092 0.0295
-11.500 -0.5412 0.02101 0.01516 -0.1319 0.7970 0.0298
-11.250 -0.5253 0.02032 0.01438 -0.1303 0.7866 0.0301
-11.000 -0.5076 0.01974 0.01371 -0.1288 0.7765 0.0303
-10.750 -0.4887 0.01921 0.01309 -0.1275 0.7673 0.0306
-10.500 -0.4693 0.01871 0.01249 -0.1262 0.7577 0.0309
-10.250 -0.4488 0.01825 0.01194 -0.1251 0.7487 0.0313
-10.000 -0.4280 0.01780 0.01138 -0.1239 0.7395 0.0316
-9.750 -0.4063 0.01738 0.01087 -0.1229 0.7316 0.0321
-9.500 -0.3843 0.01696 0.01035 -0.1219 0.7236 0.0325
-9.250 -0.3621 0.01656 0.00984 -0.1209 0.7163 0.0329
-9.000 -0.3392 0.01616 0.00936 -0.1200 0.7084 0.0333
-8.750 -0.3164 0.01581 0.00890 -0.1190 0.7005 0.0336
-8.500 -0.2940 0.01536 0.00839 -0.1181 0.6929 0.0340
-8.250 -0.2711 0.01499 0.00797 -0.1172 0.6849 0.0345
-8.000 -0.2474 0.01467 0.00759 -0.1164 0.6781 0.0350
-7.750 -0.2232 0.01436 0.00723 -0.1156 0.6708 0.0355
-7.500 -0.1990 0.01409 0.00689 -0.1148 0.6636 0.0361
-7.250 -0.1743 0.01383 0.00657 -0.1141 0.6565 0.0367
-7.000 -0.1497 0.01358 0.00625 -0.1134 0.6486 0.0374
-6.750 -0.1250 0.01336 0.00595 -0.1127 0.6413 0.0382
-6.500 -0.1003 0.01308 0.00565 -0.1120 0.6338 0.0390
-6.250 -0.0756 0.01287 0.00539 -0.1113 0.6268 0.0400
-6.000 -0.0501 0.01267 0.00514 -0.1107 0.6203 0.0410
-5.750 -0.0247 0.01249 0.00491 -0.1101 0.6129 0.0421
-5.500 0.0003 0.01233 0.00469 -0.1094 0.6060 0.0434
-5.250 0.0260 0.01214 0.00449 -0.1089 0.5991 0.0451
-5.000 0.0515 0.01200 0.00430 -0.1082 0.5924 0.0472
-4.750 0.0771 0.01186 0.00414 -0.1077 0.5867 0.0500
-4.500 0.1032 0.01173 0.00399 -0.1072 0.5805 0.0536
-4.250 0.1288 0.01163 0.00388 -0.1066 0.5742 0.0581
-4.000 0.1547 0.01155 0.00378 -0.1061 0.5687 0.0626
-3.750 0.1812 0.01147 0.00368 -0.1057 0.5632 0.0675
-3.500 0.2073 0.01141 0.00361 -0.1052 0.5580 0.0720
-3.250 0.2332 0.01139 0.00353 -0.1046 0.5530 0.0759
-3.000 0.2598 0.01131 0.00348 -0.1043 0.5482 0.0802
-2.750 0.2863 0.01129 0.00342 -0.1039 0.5432 0.0842
-2.500 0.3123 0.01124 0.00336 -0.1034 0.5387 0.0879
-2.250 0.3382 0.01122 0.00331 -0.1029 0.5346 0.0917
-2.000 0.3651 0.01118 0.00326 -0.1025 0.5304 0.0948
-1.750 0.3912 0.01113 0.00321 -0.1021 0.5260 0.0987
-1.500 0.4172 0.01110 0.00318 -0.1016 0.5219 0.1039
-1.250 0.4430 0.01110 0.00315 -0.1011 0.5181 0.1093
-1.000 0.4691 0.01104 0.00311 -0.1007 0.5146 0.1157
-0.750 0.4955 0.01101 0.00309 -0.1003 0.5105 0.1207
-0.500 0.5213 0.01097 0.00307 -0.0998 0.5063 0.1293
-0.250 0.5463 0.01094 0.00306 -0.0992 0.5017 0.1442
0.000 0.5709 0.01083 0.00307 -0.0986 0.4972 0.1868
0.250 0.5958 0.01070 0.00309 -0.0980 0.4921 0.2372
0.500 0.6200 0.01060 0.00313 -0.0973 0.4871 0.2890
0.750 0.6432 0.01054 0.00317 -0.0964 0.4817 0.3437
1.000 0.6674 0.01040 0.00323 -0.0957 0.4768 0.4102
1.250 0.6907 0.01028 0.00330 -0.0948 0.4714 0.4806
1.500 0.7128 0.01017 0.00338 -0.0937 0.4671 0.5553
1.750 0.7342 0.01007 0.00347 -0.0924 0.4635 0.6252
2.000 0.7544 0.00985 0.00357 -0.0908 0.4602 0.7240
2.250 0.7890 0.00959 0.00382 -0.0918 0.4558 0.9185
2.500 0.8518 0.00981 0.00400 -0.0992 0.4498 0.9726
2.750 0.8930 0.00998 0.00411 -0.1021 0.4449 0.9843
3.000 0.9351 0.01013 0.00424 -0.1052 0.4397 0.9981
3.250 0.9617 0.01026 0.00434 -0.1051 0.4349 1.0000
3.500 0.9803 0.01041 0.00444 -0.1033 0.4300 1.0000
3.750 1.0005 0.01052 0.00453 -0.1018 0.4246 1.0000
4.000 1.0193 0.01066 0.00464 -0.1000 0.4175 1.0000
4.250 1.0368 0.01084 0.00476 -0.0980 0.4109 1.0000
4.500 1.0563 0.01098 0.00489 -0.0964 0.4044 1.0000
4.750 1.0722 0.01116 0.00503 -0.0941 0.3974 1.0000
5.000 1.0885 0.01132 0.00517 -0.0919 0.3914 1.0000
5.250 1.1047 0.01150 0.00532 -0.0897 0.3844 1.0000
5.500 1.1196 0.01173 0.00551 -0.0873 0.3779 1.0000
5.750 1.1373 0.01193 0.00569 -0.0855 0.3708 1.0000
6.000 1.1520 0.01221 0.00593 -0.0831 0.3631 1.0000
6.250 1.1693 0.01246 0.00616 -0.0814 0.3548 1.0000
6.500 1.1834 0.01281 0.00646 -0.0791 0.3462 1.0000
6.750 1.1994 0.01314 0.00676 -0.0772 0.3362 1.0000
7.000 1.2135 0.01354 0.00712 -0.0750 0.3271 1.0000
7.250 1.2269 0.01399 0.00753 -0.0729 0.3167 1.0000
7.500 1.2413 0.01443 0.00794 -0.0709 0.3075 1.0000
8.000 1.2686 0.01545 0.00891 -0.0670 0.2909 1.0000
8.500 1.2953 0.01661 0.01003 -0.0634 0.2768 1.0000
8.750 1.3080 0.01726 0.01068 -0.0616 0.2701 1.0000
9.000 1.3209 0.01794 0.01135 -0.0600 0.2647 1.0000
9.250 1.3345 0.01861 0.01203 -0.0585 0.2588 1.0000
9.500 1.3455 0.01946 0.01286 -0.0568 0.2528 1.0000
9.750 1.3578 0.02026 0.01367 -0.0553 0.2479 1.0000
10.000 1.3712 0.02105 0.01448 -0.0540 0.2434 1.0000
10.250 1.3823 0.02199 0.01543 -0.0526 0.2387 1.0000
10.500 1.3924 0.02303 0.01647 -0.0512 0.2348 1.0000
10.750 1.4053 0.02394 0.01742 -0.0501 0.2312 1.0000
11.000 1.4178 0.02489 0.01840 -0.0491 0.2277 1.0000
11.250 1.4286 0.02598 0.01952 -0.0480 0.2242 1.0000
11.500 1.4374 0.02725 0.02080 -0.0468 0.2207 1.0000
11.750 1.4465 0.02853 0.02211 -0.0457 0.2176 1.0000
12.000 1.4590 0.02959 0.02323 -0.0449 0.2152 1.0000
12.250 1.4699 0.03078 0.02447 -0.0441 0.2124 1.0000
12.500 1.4795 0.03210 0.02584 -0.0432 0.2095 1.0000
12.750 1.4873 0.03361 0.02738 -0.0423 0.2065 1.0000
13.000 1.4935 0.03528 0.02907 -0.0414 0.2035 1.0000
13.250 1.5027 0.03670 0.03055 -0.0407 0.2009 1.0000
13.500 1.5127 0.03810 0.03202 -0.0401 0.1984 1.0000
13.750 1.5213 0.03963 0.03361 -0.0395 0.1953 1.0000
14.000 1.5249 0.04167 0.03566 -0.0388 0.1906 1.0000
14.250 1.5283 0.04374 0.03777 -0.0381 0.1870 1.0000
14.500 1.5365 0.04539 0.03947 -0.0377 0.1822 1.0000
14.750 1.5390 0.04762 0.04174 -0.0371 0.1767 1.0000
15.000 1.5408 0.04995 0.04410 -0.0366 0.1717 1.0000
15.250 1.5423 0.05236 0.04653 -0.0362 0.1640 1.0000
15.500 1.5408 0.05511 0.04929 -0.0358 0.1575 1.0000
15.750 1.5379 0.05803 0.05222 -0.0354 0.1498 1.0000
16.000 1.5334 0.06117 0.05536 -0.0351 0.1423 1.0000
16.250 1.5283 0.06442 0.05862 -0.0349 0.1366 1.0000
16.500 1.5248 0.06753 0.06176 -0.0347 0.1317 1.0000
16.750 1.5223 0.07053 0.06479 -0.0346 0.1280 1.0000
17.000 1.5196 0.07362 0.06792 -0.0346 0.1252 1.0000
17.250 1.5179 0.07661 0.07096 -0.0346 0.1225 1.0000
17.500 1.5189 0.07929 0.07371 -0.0347 0.1208 1.0000
17.750 1.5187 0.08212 0.07660 -0.0348 0.1186 1.0000
18.000 1.5151 0.08540 0.07992 -0.0350 0.1158 1.0000
18.250 1.5113 0.08876 0.08332 -0.0353 0.1136 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 527 AIRFOIL (goe527-il)