GOE 287 AIRFOIL (goe287-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 287 AIRFOIL (goe287-il) Reynolds number: 500,000 Max Cl/Cd: 106.35 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe287-il-500000.txt Download as CSV file: xf-goe287-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 287 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3740 0.10102 0.09908 -0.0107 0.9609 0.0190
-8.250 -0.3735 0.09793 0.09595 -0.0118 0.9204 0.0191
-8.000 -0.3734 0.09481 0.09273 -0.0128 0.8788 0.0191
-7.750 -0.3679 0.09116 0.08897 -0.0157 0.8517 0.0191
-7.500 -0.3650 0.08666 0.08438 -0.0147 0.8343 0.0194
-7.250 -0.3552 0.08376 0.08140 -0.0149 0.8188 0.0197
-7.000 -0.3435 0.08067 0.07824 -0.0170 0.8050 0.0201
-6.750 -0.3298 0.07733 0.07481 -0.0203 0.7921 0.0206
-6.500 -0.3139 0.07371 0.07112 -0.0244 0.7797 0.0213
-6.250 -0.2950 0.06978 0.06711 -0.0293 0.7675 0.0225
-6.000 -0.2542 0.06375 0.06088 -0.0422 0.7553 0.0242
-5.750 -0.2286 0.05898 0.05595 -0.0470 0.7428 0.0243
-5.500 -0.2129 0.05229 0.04913 -0.0507 0.7312 0.0248
-5.250 -0.1944 0.04978 0.04653 -0.0516 0.7171 0.0253
-5.000 -0.1725 0.04735 0.04398 -0.0530 0.7030 0.0260
-4.750 -0.1474 0.04469 0.04117 -0.0550 0.6897 0.0273
-4.500 -0.1063 0.04149 0.03758 -0.0584 0.6779 0.0311
-4.250 -0.0762 0.03816 0.03393 -0.0601 0.6669 0.0314
-4.000 -0.0544 0.03193 0.02746 -0.0626 0.6581 0.0326
-3.750 -0.0295 0.03041 0.02581 -0.0634 0.6488 0.0336
-3.500 -0.0027 0.02888 0.02414 -0.0641 0.6394 0.0354
-3.250 0.0283 0.02715 0.02211 -0.0646 0.6316 0.0393
-3.000 0.0599 0.02391 0.01831 -0.0654 0.6247 0.0419
-2.750 0.0866 0.02218 0.01649 -0.0661 0.6187 0.0433
-2.500 0.1143 0.02117 0.01540 -0.0666 0.6123 0.0453
-2.250 0.1435 0.02029 0.01431 -0.0669 0.6065 0.0494
-2.000 0.1739 0.01871 0.01234 -0.0672 0.6012 0.0549
-1.750 0.2020 0.01771 0.01131 -0.0677 0.5956 0.0572
-1.500 0.2309 0.01709 0.01053 -0.0679 0.5908 0.0613
-0.750 0.3224 0.01269 0.00527 -0.0676 0.5778 0.0428
-0.500 0.3513 0.01221 0.00470 -0.0677 0.5740 0.0423
-0.250 0.3803 0.01172 0.00419 -0.0678 0.5701 0.0423
0.000 0.4091 0.01130 0.00375 -0.0679 0.5660 0.0425
0.250 0.4378 0.01096 0.00337 -0.0681 0.5622 0.0435
0.500 0.4666 0.01073 0.00312 -0.0683 0.5582 0.0448
0.750 0.4957 0.01049 0.00289 -0.0685 0.5537 0.0454
1.000 0.5247 0.01033 0.00272 -0.0687 0.5494 0.0469
1.250 0.5534 0.01027 0.00260 -0.0689 0.5452 0.0491
1.500 0.5825 0.01015 0.00251 -0.0691 0.5405 0.0516
1.750 0.6115 0.01007 0.00242 -0.0693 0.5357 0.0570
2.000 0.6360 0.00829 0.00260 -0.0695 0.5314 0.8198
2.250 0.6629 0.00797 0.00252 -0.0688 0.5271 1.0000
2.500 0.6917 0.00803 0.00254 -0.0690 0.5222 1.0000
2.750 0.7203 0.00813 0.00258 -0.0692 0.5173 1.0000
3.000 0.7490 0.00820 0.00264 -0.0694 0.5117 1.0000
3.250 0.7776 0.00826 0.00268 -0.0696 0.5046 1.0000
3.500 0.8061 0.00834 0.00274 -0.0697 0.4972 1.0000
3.750 0.8347 0.00841 0.00280 -0.0699 0.4884 1.0000
4.000 0.8632 0.00848 0.00287 -0.0701 0.4791 1.0000
4.250 0.8916 0.00856 0.00294 -0.0703 0.4673 1.0000
4.500 0.9199 0.00865 0.00303 -0.0705 0.4483 1.0000
4.750 0.9473 0.00891 0.00310 -0.0706 0.3989 1.0000
5.000 0.9716 0.00982 0.00354 -0.0707 0.3211 1.0000
5.250 0.9977 0.01036 0.00392 -0.0708 0.2951 1.0000
5.500 1.0242 0.01077 0.00425 -0.0709 0.2780 1.0000
5.750 1.0509 0.01111 0.00456 -0.0710 0.2658 1.0000
6.000 1.0774 0.01147 0.00486 -0.0710 0.2528 1.0000
6.250 1.1038 0.01182 0.00517 -0.0711 0.2396 1.0000
6.500 1.1299 0.01219 0.00549 -0.0711 0.2205 1.0000
6.750 1.1497 0.01351 0.00619 -0.0707 0.1225 1.0000
7.000 1.1639 0.01559 0.00777 -0.0695 0.0293 1.0000
7.250 1.1869 0.01630 0.00855 -0.0690 0.0249 1.0000
7.500 1.2095 0.01702 0.00940 -0.0684 0.0227 1.0000
7.750 1.2320 0.01767 0.01012 -0.0679 0.0210 1.0000
8.000 1.2530 0.01845 0.01097 -0.0672 0.0195 1.0000
8.250 1.2716 0.01945 0.01205 -0.0663 0.0183 1.0000
8.500 1.2838 0.02102 0.01374 -0.0646 0.0173 1.0000
8.750 1.2961 0.02239 0.01523 -0.0629 0.0168 1.0000
9.000 1.3114 0.02335 0.01628 -0.0616 0.0163 1.0000
9.250 1.3242 0.02442 0.01742 -0.0600 0.0156 1.0000
9.500 1.3317 0.02561 0.01869 -0.0577 0.0150 1.0000
9.750 1.3370 0.02709 0.02025 -0.0557 0.0145 1.0000
10.000 1.3417 0.02882 0.02206 -0.0540 0.0141 1.0000
10.250 1.3461 0.03076 0.02408 -0.0527 0.0138 1.0000
10.500 1.3499 0.03287 0.02626 -0.0515 0.0134 1.0000
10.750 1.3527 0.03515 0.02859 -0.0504 0.0132 1.0000
11.000 1.3547 0.03754 0.03102 -0.0489 0.0129 1.0000
11.250 1.3590 0.03979 0.03328 -0.0467 0.0127 1.0000
11.500 1.3785 0.04155 0.03499 -0.0435 0.0124 1.0000
11.750 1.3827 0.04343 0.03704 -0.0430 0.0122 1.0000
12.000 1.3889 0.04537 0.03914 -0.0423 0.0120 1.0000
12.250 1.3958 0.04738 0.04128 -0.0416 0.0117 1.0000
12.500 1.4035 0.04947 0.04349 -0.0406 0.0114 1.0000
12.750 1.4130 0.05163 0.04578 -0.0392 0.0113 1.0000
13.000 1.4220 0.05404 0.04833 -0.0379 0.0112 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 287 AIRFOIL (goe287-il)