GOE 572 AIRFOIL (goe572-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 572 AIRFOIL (goe572-il) Reynolds number: 1,000,000 Max Cl/Cd: 103.81 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe572-il-1000000.txt Download as CSV file: xf-goe572-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 572 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.2484 0.02720 0.02320 -0.1525 0.5937 0.0155
-11.000 -0.2674 0.02477 0.02047 -0.1464 0.5865 0.0157
-10.750 -0.2621 0.02372 0.01925 -0.1431 0.5793 0.0160
-10.500 -0.2625 0.02230 0.01759 -0.1389 0.5722 0.0162
-10.250 -0.2545 0.02137 0.01648 -0.1357 0.5652 0.0164
-10.000 -0.2500 0.02015 0.01503 -0.1319 0.5589 0.0167
-9.750 -0.2424 0.01922 0.01394 -0.1285 0.5524 0.0171
-9.500 -0.2249 0.01885 0.01352 -0.1265 0.5471 0.0174
-9.250 -0.2055 0.01870 0.01333 -0.1249 0.5408 0.0177
-9.000 -0.1819 0.01890 0.01352 -0.1239 0.5352 0.0181
-8.750 -0.1670 0.01823 0.01273 -0.1214 0.5302 0.0185
-8.500 -0.1510 0.01769 0.01205 -0.1192 0.5249 0.0189
-8.250 -0.1337 0.01732 0.01155 -0.1171 0.5196 0.0194
-8.000 -0.1137 0.01699 0.01111 -0.1154 0.5157 0.0198
-7.750 -0.0990 0.01621 0.01017 -0.1129 0.5114 0.0203
-7.500 -0.0791 0.01609 0.01002 -0.1113 0.5059 0.0209
-7.250 -0.0588 0.01590 0.00980 -0.1097 0.5019 0.0214
-7.000 -0.0371 0.01575 0.00961 -0.1084 0.4981 0.0219
-6.750 -0.0160 0.01559 0.00938 -0.1069 0.4939 0.0226
-6.500 0.0029 0.01524 0.00891 -0.1050 0.4891 0.0231
-6.250 0.0238 0.01503 0.00859 -0.1034 0.4855 0.0237
-6.000 0.0428 0.01442 0.00787 -0.1016 0.4817 0.0243
-5.750 0.0629 0.01406 0.00749 -0.1000 0.4779 0.0250
-5.500 0.0842 0.01404 0.00743 -0.0985 0.4737 0.0257
-5.250 0.1055 0.01397 0.00731 -0.0971 0.4697 0.0264
-5.000 0.1274 0.01369 0.00699 -0.0958 0.4664 0.0271
-4.750 0.1490 0.01349 0.00671 -0.0943 0.4624 0.0278
-4.500 0.1699 0.01339 0.00653 -0.0927 0.4584 0.0283
-4.250 0.1872 0.01286 0.00591 -0.0905 0.4543 0.0293
-4.000 0.2079 0.01261 0.00567 -0.0890 0.4516 0.0301
-3.750 0.2288 0.01252 0.00557 -0.0874 0.4479 0.0309
-3.500 0.2485 0.01236 0.00535 -0.0856 0.4438 0.0318
-3.250 0.2676 0.01225 0.00518 -0.0837 0.4397 0.0325
-3.000 0.2891 0.01213 0.00503 -0.0823 0.4367 0.0333
-2.750 0.3099 0.01190 0.00476 -0.0807 0.4330 0.0341
-2.500 0.3286 0.01160 0.00442 -0.0788 0.4292 0.0350
-2.250 0.3482 0.01151 0.00430 -0.0770 0.4252 0.0361
-2.000 0.3687 0.01142 0.00418 -0.0754 0.4215 0.0369
-1.750 0.3903 0.01131 0.00406 -0.0741 0.4182 0.0379
-1.500 0.4113 0.01127 0.00399 -0.0726 0.4147 0.0390
-1.250 0.4315 0.01126 0.00394 -0.0710 0.4109 0.0397
-1.000 0.4495 0.01114 0.00377 -0.0689 0.4068 0.0412
-0.750 0.4717 0.01106 0.00369 -0.0677 0.4037 0.0430
-0.500 0.4933 0.01104 0.00366 -0.0664 0.4000 0.0443
-0.250 0.5138 0.01107 0.00365 -0.0649 0.3957 0.0461
0.000 0.5334 0.01110 0.00363 -0.0633 0.3912 0.0486
0.250 0.5558 0.01105 0.00361 -0.0621 0.3886 0.0528
0.500 0.5774 0.01108 0.00361 -0.0609 0.3851 0.0553
0.750 0.5978 0.01108 0.00362 -0.0595 0.3814 0.0653
1.000 0.6171 0.01112 0.00369 -0.0579 0.3773 0.0841
1.250 0.6393 0.01105 0.00375 -0.0569 0.3742 0.1270
1.500 0.6622 0.01086 0.00385 -0.0562 0.3710 0.2382
1.750 0.6894 0.01044 0.00403 -0.0567 0.3675 0.4735
2.000 1.0875 0.01066 0.00515 -0.1383 0.3521 1.0000
2.250 1.1035 0.01079 0.00524 -0.1361 0.3489 1.0000
2.500 1.1172 0.01094 0.00535 -0.1335 0.3455 1.0000
2.750 1.1333 0.01101 0.00542 -0.1312 0.3434 1.0000
3.000 1.1488 0.01111 0.00550 -0.1289 0.3410 1.0000
3.250 1.1640 0.01123 0.00560 -0.1266 0.3380 1.0000
3.500 1.1789 0.01139 0.00573 -0.1242 0.3348 1.0000
3.750 1.1932 0.01159 0.00589 -0.1217 0.3315 1.0000
4.000 1.2107 0.01171 0.00601 -0.1199 0.3294 1.0000
4.250 1.2285 0.01184 0.00614 -0.1181 0.3268 1.0000
4.500 1.2457 0.01200 0.00629 -0.1163 0.3246 1.0000
4.750 1.2620 0.01220 0.00647 -0.1144 0.3215 1.0000
5.000 1.2772 0.01244 0.00668 -0.1122 0.3187 1.0000
5.250 1.2915 0.01270 0.00693 -0.1099 0.3156 1.0000
5.500 1.3104 0.01285 0.00709 -0.1086 0.3138 1.0000
5.750 1.3283 0.01304 0.00729 -0.1070 0.3119 1.0000
6.000 1.3450 0.01327 0.00752 -0.1053 0.3095 1.0000
6.250 1.3609 0.01354 0.00778 -0.1035 0.3070 1.0000
6.500 1.3739 0.01391 0.00812 -0.1012 0.3028 1.0000
6.750 1.3883 0.01425 0.00847 -0.0992 0.3010 1.0000
7.000 1.4076 0.01445 0.00869 -0.0981 0.2986 1.0000
7.250 1.4242 0.01475 0.00901 -0.0966 0.2969 1.0000
7.500 1.4399 0.01509 0.00936 -0.0950 0.2940 1.0000
7.750 1.4534 0.01553 0.00979 -0.0931 0.2909 1.0000
8.000 1.4643 0.01609 0.01034 -0.0908 0.2863 1.0000
8.250 1.4815 0.01643 0.01071 -0.0897 0.2845 1.0000
8.500 1.4971 0.01685 0.01115 -0.0883 0.2823 1.0000
8.750 1.5100 0.01741 0.01171 -0.0866 0.2783 1.0000
9.000 1.5183 0.01821 0.01249 -0.0844 0.2733 1.0000
9.250 1.5332 0.01874 0.01304 -0.0832 0.2704 1.0000
9.500 1.5466 0.01936 0.01369 -0.0818 0.2670 1.0000
9.750 1.5564 0.02020 0.01452 -0.0801 0.2631 1.0000
10.000 1.5627 0.02128 0.01560 -0.0780 0.2579 1.0000
10.250 1.5752 0.02205 0.01640 -0.0768 0.2541 1.0000
10.500 1.5786 0.02342 0.01775 -0.0748 0.2472 1.0000
10.750 1.5840 0.02473 0.01907 -0.0731 0.2406 1.0000
11.000 1.5840 0.02650 0.02082 -0.0711 0.2336 1.0000
11.250 1.5839 0.02837 0.02268 -0.0694 0.2254 1.0000
11.500 1.5780 0.03078 0.02509 -0.0674 0.2182 1.0000
11.750 1.5689 0.03359 0.02788 -0.0654 0.2080 1.0000
12.000 1.5474 0.03760 0.03186 -0.0630 0.1947 1.0000
12.250 1.5314 0.04133 0.03559 -0.0612 0.1841 1.0000
12.500 1.4819 0.04845 0.04265 -0.0584 0.1600 1.0000
12.750 1.4671 0.05244 0.04668 -0.0573 0.1520 1.0000
13.000 1.4336 0.05848 0.05272 -0.0557 0.1351 1.0000
13.250 1.3962 0.06521 0.05949 -0.0544 0.1178 1.0000
13.500 1.3786 0.06990 0.06424 -0.0538 0.1102 1.0000
13.750 1.3532 0.07562 0.07001 -0.0531 0.0991 1.0000
14.000 1.3223 0.08221 0.07664 -0.0527 0.0839 1.0000
14.250 1.3012 0.08772 0.08217 -0.0524 0.0724 1.0000
14.500 1.2771 0.09379 0.08825 -0.0524 0.0593 1.0000
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Polar data table (+)
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