GOE 565 AIRFOIL (goe565-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: GOE 565 AIRFOIL (goe565-il) Reynolds number: 100,000 Max Cl/Cd: 55.04 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe565-il-100000-n5.txt Download as CSV file: xf-goe565-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 565 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4405 0.08904 0.08390 -0.0342 1.0000 0.0333
-8.750 -0.4462 0.08500 0.07995 -0.0354 1.0000 0.0333
-8.500 -0.4547 0.08105 0.07610 -0.0363 1.0000 0.0334
-8.250 -0.4662 0.07741 0.07258 -0.0368 1.0000 0.0333
-8.000 -0.4773 0.07294 0.06820 -0.0387 1.0000 0.0333
-7.750 -0.4866 0.06757 0.06287 -0.0419 1.0000 0.0333
-7.500 -0.4957 0.06101 0.05628 -0.0455 1.0000 0.0334
-7.250 -0.5044 0.05365 0.04874 -0.0486 1.0000 0.0338
-7.000 -0.5119 0.04547 0.04004 -0.0509 1.0000 0.0351
-6.750 -0.5083 0.04038 0.03448 -0.0511 1.0000 0.0368
-6.500 -0.4949 0.03899 0.03302 -0.0501 1.0000 0.0383
-6.250 -0.4820 0.03622 0.02993 -0.0494 1.0000 0.0401
-6.000 -0.4672 0.03313 0.02629 -0.0489 1.0000 0.0438
-5.750 -0.4506 0.03008 0.02256 -0.0482 1.0000 0.0472
-5.500 -0.4324 0.02861 0.02096 -0.0474 1.0000 0.0496
-5.250 -0.4043 0.02738 0.01944 -0.0483 0.9974 0.0544
-5.000 -0.3719 0.02552 0.01694 -0.0496 0.9939 0.0583
-4.750 -0.3413 0.02415 0.01543 -0.0509 0.9900 0.0624
-4.500 -0.3083 0.02302 0.01404 -0.0523 0.9866 0.0653
-4.250 -0.2769 0.02193 0.01271 -0.0532 0.9824 0.0678
-4.000 -0.2438 0.02114 0.01164 -0.0543 0.9785 0.0715
-3.750 -0.2123 0.02033 0.01080 -0.0554 0.9744 0.0746
-3.500 -0.1803 0.01966 0.01008 -0.0564 0.9698 0.0771
-3.250 -0.1471 0.01907 0.00943 -0.0576 0.9658 0.0802
-3.000 -0.1162 0.01858 0.00883 -0.0583 0.9602 0.0834
-2.750 -0.0807 0.01809 0.00836 -0.0600 0.9562 0.0885
-2.500 -0.0496 0.01774 0.00795 -0.0607 0.9487 0.0963
-2.250 -0.0141 0.01725 0.00746 -0.0623 0.9432 0.1052
-2.000 0.0179 0.01679 0.00703 -0.0631 0.9360 0.1192
-1.750 0.0533 0.01620 0.00664 -0.0648 0.9313 0.1567
-1.500 0.0821 0.01555 0.00656 -0.0654 0.9241 0.2984
-1.250 0.1181 0.01504 0.00645 -0.0671 0.9199 0.4121
-1.000 0.1443 0.01456 0.00649 -0.0667 0.9121 0.5515
-0.750 0.1755 0.01395 0.00643 -0.0667 0.9071 0.7172
-0.500 0.2407 0.01343 0.00628 -0.0736 0.9063 1.0000
-0.250 0.2698 0.01344 0.00617 -0.0738 0.8969 1.0000
0.000 0.3081 0.01334 0.00595 -0.0756 0.8900 1.0000
0.250 0.3369 0.01328 0.00581 -0.0754 0.8778 1.0000
0.500 0.3667 0.01323 0.00568 -0.0755 0.8661 1.0000
0.750 0.3975 0.01317 0.00557 -0.0757 0.8556 1.0000
1.000 0.4298 0.01309 0.00544 -0.0762 0.8455 1.0000
1.250 0.4575 0.01305 0.00537 -0.0757 0.8316 1.0000
1.500 0.4858 0.01300 0.00528 -0.0754 0.8167 1.0000
1.750 0.5150 0.01293 0.00518 -0.0751 0.8005 1.0000
2.000 0.5422 0.01291 0.00513 -0.0745 0.7819 1.0000
2.250 0.5701 0.01290 0.00510 -0.0740 0.7628 1.0000
2.500 0.5987 0.01290 0.00507 -0.0736 0.7431 1.0000
2.750 0.6255 0.01295 0.00508 -0.0729 0.7196 1.0000
3.000 0.6522 0.01303 0.00511 -0.0721 0.6936 1.0000
3.250 0.6776 0.01317 0.00522 -0.0712 0.6662 1.0000
3.500 0.7023 0.01334 0.00535 -0.0703 0.6370 1.0000
3.750 0.7270 0.01354 0.00550 -0.0693 0.6058 1.0000
4.000 0.7513 0.01379 0.00568 -0.0683 0.5729 1.0000
4.250 0.7748 0.01410 0.00590 -0.0672 0.5391 1.0000
4.500 0.7976 0.01449 0.00616 -0.0659 0.5039 1.0000
4.750 0.8189 0.01496 0.00648 -0.0645 0.4624 1.0000
5.000 0.8391 0.01550 0.00686 -0.0630 0.4157 1.0000
5.250 0.8584 0.01611 0.00725 -0.0614 0.3677 1.0000
5.500 0.8781 0.01675 0.00771 -0.0600 0.3264 1.0000
5.750 0.8981 0.01738 0.00822 -0.0586 0.2905 1.0000
6.000 0.9181 0.01803 0.00878 -0.0574 0.2539 1.0000
6.250 0.9376 0.01874 0.00934 -0.0561 0.2188 1.0000
6.500 0.9567 0.01950 0.00997 -0.0548 0.1893 1.0000
6.750 0.9764 0.02024 0.01068 -0.0536 0.1676 1.0000
7.000 0.9958 0.02102 0.01146 -0.0523 0.1422 1.0000
7.250 1.0148 0.02189 0.01228 -0.0509 0.1026 1.0000
7.500 1.0299 0.02315 0.01329 -0.0492 0.0774 1.0000
7.750 1.0459 0.02431 0.01447 -0.0474 0.0654 1.0000
8.000 1.0615 0.02547 0.01566 -0.0457 0.0575 1.0000
8.250 1.0766 0.02664 0.01689 -0.0439 0.0514 1.0000
8.500 1.0905 0.02792 0.01826 -0.0419 0.0473 1.0000
8.750 1.1045 0.02915 0.01961 -0.0401 0.0429 1.0000
9.000 1.1149 0.03076 0.02129 -0.0379 0.0399 1.0000
9.250 1.1277 0.03234 0.02302 -0.0359 0.0373 1.0000
9.500 1.1407 0.03383 0.02469 -0.0340 0.0344 1.0000
9.750 1.1519 0.03550 0.02647 -0.0320 0.0322 1.0000
10.000 1.1628 0.03751 0.02854 -0.0302 0.0305 1.0000
10.250 1.1747 0.03995 0.03114 -0.0286 0.0287 1.0000
10.500 1.1838 0.04176 0.03326 -0.0265 0.0272 1.0000
10.750 1.1913 0.04390 0.03568 -0.0244 0.0258 1.0000
11.000 1.1961 0.04637 0.03842 -0.0222 0.0248 1.0000
11.250 1.1982 0.04893 0.04124 -0.0201 0.0241 1.0000
11.500 1.1969 0.05157 0.04417 -0.0180 0.0234 1.0000
11.750 1.1934 0.05428 0.04709 -0.0162 0.0227 1.0000
12.000 1.1887 0.05709 0.05008 -0.0147 0.0221 1.0000
12.250 1.1818 0.06035 0.05351 -0.0137 0.0217 1.0000
12.500 1.1729 0.06411 0.05742 -0.0131 0.0212 1.0000
13.000 1.1427 0.07342 0.06718 -0.0141 0.0208 1.0000
13.250 1.1252 0.07877 0.07278 -0.0159 0.0207 1.0000
13.500 1.1065 0.08491 0.07914 -0.0188 0.0207 1.0000
13.750 1.0877 0.09169 0.08611 -0.0224 0.0208 1.0000
14.000 1.0678 0.09926 0.09386 -0.0271 0.0209 1.0000
14.250 0.9677 0.13195 0.12711 -0.0503 0.0247 1.0000
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Polar data table (+)
Polar graphs
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