GOE 563 AIRFOIL (goe563-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 563 AIRFOIL (goe563-il) Reynolds number: 200,000 Max Cl/Cd: 73.32 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe563-il-200000.txt Download as CSV file: xf-goe563-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 563 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.4500 0.10745 0.10359 -0.0262 1.0000 0.0571
-10.000 -0.4639 0.10352 0.09973 -0.0320 1.0000 0.0594
-9.750 -0.4776 0.09926 0.09555 -0.0373 1.0000 0.0599
-9.500 -0.4770 0.09383 0.09016 -0.0369 1.0000 0.0606
-9.250 -0.4587 0.09188 0.08818 -0.0327 1.0000 0.0622
-9.000 -0.4502 0.08978 0.08609 -0.0316 1.0000 0.0647
-8.750 -0.4509 0.08673 0.08308 -0.0323 1.0000 0.0669
-8.500 -0.4620 0.08302 0.07944 -0.0348 1.0000 0.0698
-8.250 -0.4074 0.06737 0.06408 -0.0385 1.0000 0.0740
-8.000 -0.4024 0.06512 0.06185 -0.0362 1.0000 0.0755
-7.750 -0.5358 0.07041 0.06668 -0.0437 1.0000 0.0720
-7.500 -0.5279 0.06547 0.06197 -0.0411 1.0000 0.0736
-7.250 -0.5215 0.06393 0.06049 -0.0380 1.0000 0.0749
-7.000 -0.5205 0.06180 0.05835 -0.0360 1.0000 0.0765
-6.750 -0.5213 0.05904 0.05554 -0.0348 1.0000 0.0787
-6.500 -0.5344 0.05637 0.05207 -0.0356 1.0000 0.0851
-6.250 -0.5306 0.04991 0.04584 -0.0347 1.0000 0.0868
-6.000 -0.5203 0.04791 0.04390 -0.0329 1.0000 0.0885
-5.750 -0.5222 0.03377 0.02825 -0.0317 1.0000 0.0593
-5.500 -0.5094 0.02996 0.02412 -0.0303 1.0000 0.0581
-5.250 -0.4932 0.02724 0.02095 -0.0289 1.0000 0.0584
-5.000 -0.4700 0.02542 0.01868 -0.0283 0.9990 0.0595
-4.750 -0.4360 0.02344 0.01628 -0.0298 0.9958 0.0600
-4.500 -0.4043 0.02088 0.01344 -0.0311 0.9932 0.0619
-4.250 -0.3705 0.01996 0.01242 -0.0326 0.9891 0.0644
-4.000 -0.3344 0.01897 0.01127 -0.0343 0.9855 0.0662
-3.750 -0.2981 0.01814 0.01029 -0.0361 0.9821 0.0686
-3.500 -0.2648 0.01759 0.00959 -0.0371 0.9768 0.0716
-3.250 -0.2291 0.01667 0.00870 -0.0389 0.9729 0.0756
-3.000 -0.1904 0.01618 0.00823 -0.0411 0.9697 0.0806
-2.750 -0.1611 0.01566 0.00767 -0.0415 0.9630 0.0866
-2.500 -0.1237 0.01522 0.00728 -0.0434 0.9587 0.0999
-2.250 -0.0851 0.01450 0.00678 -0.0459 0.9557 0.1333
-2.000 -0.0592 0.01410 0.00653 -0.0456 0.9478 0.1702
-1.750 -0.0219 0.01364 0.00630 -0.0477 0.9436 0.2213
-1.500 0.0072 0.01260 0.00626 -0.0486 0.9388 0.4526
-1.250 0.0268 0.01169 0.00634 -0.0466 0.9315 0.6937
-1.000 0.1455 0.01128 0.00650 -0.0642 0.9444 1.0000
-0.750 0.1956 0.01115 0.00625 -0.0686 0.9415 1.0000
-0.500 0.2319 0.01099 0.00600 -0.0701 0.9321 1.0000
-0.250 0.2861 0.01060 0.00553 -0.0750 0.9273 1.0000
0.000 0.3190 0.01038 0.00526 -0.0756 0.9166 1.0000
0.250 0.3608 0.01010 0.00492 -0.0780 0.9110 1.0000
0.500 0.3854 0.01001 0.00480 -0.0771 0.8994 1.0000
0.750 0.4127 0.00990 0.00466 -0.0766 0.8891 1.0000
1.000 0.4443 0.00971 0.00444 -0.0768 0.8807 1.0000
1.250 0.4677 0.00965 0.00436 -0.0755 0.8681 1.0000
1.500 0.4918 0.00958 0.00428 -0.0743 0.8555 1.0000
1.750 0.5164 0.00950 0.00417 -0.0731 0.8416 1.0000
2.000 0.5409 0.00940 0.00405 -0.0718 0.8257 1.0000
2.250 0.5653 0.00932 0.00393 -0.0705 0.8084 1.0000
2.500 0.5890 0.00928 0.00385 -0.0691 0.7884 1.0000
2.750 0.6122 0.00927 0.00380 -0.0676 0.7656 1.0000
3.000 0.6354 0.00931 0.00379 -0.0661 0.7411 1.0000
3.250 0.6589 0.00939 0.00379 -0.0648 0.7167 1.0000
3.500 0.6813 0.00951 0.00386 -0.0633 0.6890 1.0000
3.750 0.7034 0.00967 0.00395 -0.0618 0.6574 1.0000
4.000 0.7244 0.00988 0.00405 -0.0600 0.6176 1.0000
4.250 0.7436 0.01022 0.00416 -0.0579 0.5670 1.0000
4.500 0.7608 0.01069 0.00434 -0.0556 0.5056 1.0000
4.750 0.7777 0.01124 0.00461 -0.0533 0.4451 1.0000
5.000 0.7952 0.01180 0.00493 -0.0514 0.3934 1.0000
5.250 0.8131 0.01240 0.00530 -0.0495 0.3493 1.0000
5.500 0.8316 0.01300 0.00570 -0.0478 0.3108 1.0000
5.750 0.8513 0.01352 0.00612 -0.0464 0.2771 1.0000
6.000 0.8711 0.01405 0.00653 -0.0450 0.2433 1.0000
6.250 0.8906 0.01462 0.00696 -0.0436 0.2095 1.0000
6.500 0.9103 0.01521 0.00743 -0.0422 0.1762 1.0000
6.750 0.9283 0.01596 0.00800 -0.0407 0.1397 1.0000
7.000 0.9432 0.01708 0.00886 -0.0387 0.0985 1.0000
7.250 0.9596 0.01810 0.00975 -0.0367 0.0752 1.0000
7.500 0.9771 0.01897 0.01058 -0.0350 0.0650 1.0000
7.750 0.9925 0.02005 0.01161 -0.0330 0.0587 1.0000
8.000 1.0107 0.02089 0.01254 -0.0313 0.0538 1.0000
8.250 1.0247 0.02228 0.01387 -0.0293 0.0494 1.0000
8.500 1.0431 0.02334 0.01505 -0.0277 0.0467 1.0000
8.750 1.0614 0.02452 0.01630 -0.0262 0.0440 1.0000
9.000 1.0792 0.02584 0.01761 -0.0249 0.0414 1.0000
9.250 1.0995 0.02807 0.01993 -0.0240 0.0390 1.0000
9.500 1.1191 0.02952 0.02159 -0.0227 0.0375 1.0000
9.750 1.1383 0.03129 0.02356 -0.0215 0.0361 1.0000
10.000 1.1558 0.03302 0.02546 -0.0202 0.0347 1.0000
10.250 1.1716 0.03477 0.02731 -0.0189 0.0332 1.0000
10.500 1.1874 0.03829 0.03098 -0.0181 0.0317 1.0000
10.750 1.1935 0.04170 0.03475 -0.0158 0.0312 1.0000
11.000 1.1961 0.04392 0.03732 -0.0128 0.0309 1.0000
11.250 1.1948 0.04664 0.04037 -0.0097 0.0307 1.0000
11.500 1.1866 0.04935 0.04340 -0.0058 0.0306 1.0000
11.750 1.1750 0.05227 0.04658 -0.0021 0.0306 1.0000
12.000 1.1608 0.05550 0.05007 0.0010 0.0307 1.0000
12.250 1.1435 0.05917 0.05399 0.0034 0.0308 1.0000
12.500 1.1259 0.06340 0.05843 0.0048 0.0310 1.0000
12.750 1.1058 0.06801 0.06324 0.0054 0.0311 1.0000
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Polar data table (+)
Polar graphs
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