GOE 566 AIRFOIL (goe566-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 566 AIRFOIL (goe566-il) Reynolds number: 200,000 Max Cl/Cd: 61.69 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe566-il-200000-n5.txt Download as CSV file: xf-goe566-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 566 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4511 0.08643 0.08279 -0.0313 1.0000 0.0204
-9.000 -0.4533 0.08264 0.07903 -0.0329 1.0000 0.0207
-8.750 -0.4586 0.07860 0.07505 -0.0347 1.0000 0.0209
-8.500 -0.4677 0.07470 0.07123 -0.0364 1.0000 0.0211
-8.250 -0.4792 0.07066 0.06725 -0.0384 1.0000 0.0213
-8.000 -0.4872 0.06610 0.06270 -0.0407 1.0000 0.0218
-7.750 -0.4962 0.06148 0.05807 -0.0417 1.0000 0.0222
-7.500 -0.5064 0.05694 0.05346 -0.0413 1.0000 0.0226
-7.250 -0.5014 0.04793 0.04411 -0.0461 0.9946 0.0235
-7.000 -0.4940 0.03335 0.02830 -0.0514 0.9847 0.0257
-6.750 -0.4679 0.03203 0.02690 -0.0527 0.9793 0.0269
-6.500 -0.4381 0.03071 0.02542 -0.0543 0.9756 0.0283
-6.250 -0.4134 0.02794 0.02224 -0.0550 0.9696 0.0297
-6.000 -0.3848 0.02555 0.01937 -0.0560 0.9652 0.0320
-5.750 -0.3539 0.02383 0.01716 -0.0570 0.9617 0.0336
-5.500 -0.3285 0.02171 0.01467 -0.0571 0.9556 0.0346
-5.250 -0.2982 0.02043 0.01327 -0.0582 0.9516 0.0364
-5.000 -0.2663 0.01946 0.01214 -0.0592 0.9480 0.0380
-4.750 -0.2392 0.01846 0.01096 -0.0591 0.9416 0.0389
-4.500 -0.2080 0.01751 0.00984 -0.0598 0.9372 0.0399
-4.250 -0.1757 0.01672 0.00891 -0.0607 0.9334 0.0414
-4.000 -0.1486 0.01613 0.00820 -0.0605 0.9264 0.0429
-3.750 -0.1181 0.01544 0.00742 -0.0610 0.9216 0.0438
-3.500 -0.0893 0.01486 0.00675 -0.0611 0.9159 0.0444
-3.250 -0.0618 0.01428 0.00613 -0.0610 0.9093 0.0452
-3.000 -0.0326 0.01359 0.00543 -0.0612 0.9044 0.0467
-2.750 -0.0078 0.01318 0.00501 -0.0606 0.8964 0.0481
-2.500 0.0206 0.01282 0.00462 -0.0606 0.8906 0.0502
-2.250 0.0465 0.01255 0.00434 -0.0601 0.8828 0.0532
-2.000 0.0743 0.01229 0.00402 -0.0599 0.8761 0.0563
-1.750 0.1000 0.01203 0.00373 -0.0593 0.8670 0.0599
-1.500 0.1267 0.01174 0.00346 -0.0588 0.8573 0.0684
-1.250 0.1523 0.01138 0.00322 -0.0581 0.8452 0.1038
-1.000 0.1768 0.01101 0.00304 -0.0573 0.8308 0.1704
-0.750 0.2007 0.01058 0.00287 -0.0564 0.8160 0.2592
-0.500 0.2129 0.00919 0.00286 -0.0533 0.8025 0.6529
-0.250 0.2638 0.00856 0.00294 -0.0568 0.7883 0.9179
0.250 0.3730 0.00855 0.00271 -0.0676 0.7578 1.0000
0.500 0.3963 0.00858 0.00265 -0.0666 0.7440 1.0000
0.750 0.4198 0.00863 0.00260 -0.0656 0.7296 1.0000
1.000 0.4434 0.00869 0.00257 -0.0646 0.7144 1.0000
1.250 0.4670 0.00877 0.00255 -0.0636 0.6977 1.0000
1.500 0.4904 0.00887 0.00255 -0.0625 0.6780 1.0000
1.750 0.5135 0.00899 0.00255 -0.0614 0.6549 1.0000
2.000 0.5363 0.00914 0.00258 -0.0602 0.6290 1.0000
2.250 0.5587 0.00933 0.00262 -0.0589 0.5976 1.0000
2.500 0.5805 0.00956 0.00268 -0.0576 0.5603 1.0000
2.750 0.6013 0.00985 0.00276 -0.0561 0.5106 1.0000
3.000 0.6213 0.01023 0.00289 -0.0545 0.4609 1.0000
3.250 0.6414 0.01065 0.00307 -0.0530 0.4184 1.0000
3.500 0.6621 0.01105 0.00329 -0.0517 0.3833 1.0000
4.000 0.7058 0.01176 0.00378 -0.0495 0.3304 1.0000
4.250 0.7282 0.01209 0.00404 -0.0485 0.3113 1.0000
4.500 0.7507 0.01242 0.00432 -0.0475 0.2953 1.0000
4.750 0.7731 0.01277 0.00461 -0.0465 0.2801 1.0000
5.000 0.7961 0.01308 0.00491 -0.0457 0.2682 1.0000
5.250 0.8190 0.01341 0.00522 -0.0448 0.2569 1.0000
5.500 0.8415 0.01376 0.00556 -0.0439 0.2436 1.0000
5.750 0.8640 0.01412 0.00590 -0.0430 0.2310 1.0000
6.000 0.8863 0.01449 0.00625 -0.0421 0.2193 1.0000
6.250 0.9093 0.01481 0.00660 -0.0413 0.2068 1.0000
6.500 0.9323 0.01513 0.00698 -0.0405 0.1952 1.0000
6.750 0.9549 0.01548 0.00736 -0.0396 0.1797 1.0000
7.000 0.9769 0.01588 0.00774 -0.0387 0.1588 1.0000
7.250 0.9969 0.01646 0.00818 -0.0376 0.1258 1.0000
7.500 1.0150 0.01725 0.00881 -0.0362 0.1002 1.0000
7.750 1.0336 0.01800 0.00950 -0.0349 0.0863 1.0000
8.000 1.0531 0.01865 0.01019 -0.0337 0.0737 1.0000
8.250 1.0723 0.01933 0.01087 -0.0325 0.0569 1.0000
8.500 1.0896 0.02016 0.01160 -0.0310 0.0432 1.0000
8.750 1.1057 0.02107 0.01252 -0.0294 0.0368 1.0000
9.000 1.1215 0.02194 0.01348 -0.0277 0.0328 1.0000
9.250 1.1331 0.02312 0.01470 -0.0255 0.0293 1.0000
9.500 1.1485 0.02393 0.01566 -0.0238 0.0268 1.0000
9.750 1.1610 0.02488 0.01673 -0.0217 0.0246 1.0000
10.000 1.1696 0.02598 0.01791 -0.0191 0.0230 1.0000
10.250 1.1731 0.02740 0.01940 -0.0160 0.0216 1.0000
10.500 1.1832 0.02841 0.02057 -0.0138 0.0205 1.0000
10.750 1.1919 0.02956 0.02187 -0.0116 0.0190 1.0000
11.000 1.1997 0.03079 0.02321 -0.0096 0.0180 1.0000
11.250 1.2054 0.03222 0.02475 -0.0075 0.0173 1.0000
11.500 1.2096 0.03381 0.02644 -0.0056 0.0166 1.0000
11.750 1.2097 0.03584 0.02855 -0.0036 0.0159 1.0000
12.000 1.2114 0.03791 0.03075 -0.0019 0.0154 1.0000
12.250 1.2147 0.03989 0.03289 -0.0006 0.0151 1.0000
12.500 1.2170 0.04203 0.03521 0.0006 0.0145 1.0000
12.750 1.2185 0.04430 0.03765 0.0015 0.0138 1.0000
13.000 1.2186 0.04683 0.04033 0.0022 0.0134 1.0000
13.250 1.2182 0.04945 0.04312 0.0025 0.0129 1.0000
13.500 1.2161 0.05238 0.04618 0.0024 0.0125 1.0000
13.750 1.2124 0.05572 0.04967 0.0022 0.0124 1.0000
14.000 1.2078 0.05930 0.05340 0.0016 0.0122 1.0000
14.250 1.2020 0.06320 0.05744 0.0007 0.0120 1.0000
14.500 1.1944 0.06753 0.06190 -0.0007 0.0118 1.0000
14.750 1.1860 0.07223 0.06676 -0.0025 0.0117 1.0000
15.000 1.1753 0.07752 0.07220 -0.0047 0.0116 1.0000
15.250 1.1637 0.08325 0.07809 -0.0073 0.0114 1.0000
15.500 1.1500 0.08970 0.08470 -0.0105 0.0114 1.0000
15.750 1.1372 0.09626 0.09144 -0.0141 0.0113 1.0000
16.000 1.1232 0.10338 0.09873 -0.0181 0.0113 1.0000
16.250 1.1073 0.11119 0.10672 -0.0226 0.0113 1.0000
16.500 1.0915 0.11920 0.11489 -0.0273 0.0113 1.0000
16.750 1.0751 0.12767 0.12352 -0.0324 0.0114 1.0000
17.250 1.0306 0.14933 0.14555 -0.0456 0.0116 1.0000
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