GOE 564 AIRFOIL (goe564-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 564 AIRFOIL (goe564-il) Reynolds number: 200,000 Max Cl/Cd: 65.62 at α=3° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe564-il-200000-n5.txt Download as CSV file: xf-goe564-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 564 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 -0.4710 0.08608 0.08245 -0.0336 1.0000 0.0200
-9.000 -0.4815 0.08156 0.07800 -0.0347 1.0000 0.0202
-8.750 -0.4955 0.07738 0.07390 -0.0354 1.0000 0.0202
-8.500 -0.5164 0.07356 0.07019 -0.0353 1.0000 0.0203
-8.250 -0.5340 0.06825 0.06494 -0.0374 1.0000 0.0202
-8.000 -0.5526 0.06148 0.05816 -0.0401 1.0000 0.0203
-7.750 -0.5784 0.05080 0.04729 -0.0433 1.0000 0.0204
-7.500 -0.5925 0.02898 0.02361 -0.0517 0.9936 0.0224
-7.250 -0.5681 0.02689 0.02125 -0.0526 0.9903 0.0240
-7.000 -0.5398 0.02593 0.02016 -0.0537 0.9877 0.0255
-6.750 -0.5115 0.02415 0.01801 -0.0547 0.9851 0.0274
-6.500 -0.4843 0.02277 0.01622 -0.0551 0.9811 0.0298
-6.250 -0.4548 0.02114 0.01423 -0.0560 0.9779 0.0318
-6.000 -0.4232 0.02002 0.01297 -0.0574 0.9755 0.0339
-5.750 -0.3965 0.01939 0.01223 -0.0574 0.9708 0.0363
-5.500 -0.3659 0.01849 0.01111 -0.0582 0.9674 0.0384
-5.250 -0.3332 0.01774 0.01013 -0.0593 0.9649 0.0404
-5.000 -0.3057 0.01684 0.00911 -0.0594 0.9607 0.0422
-4.750 -0.2768 0.01610 0.00831 -0.0598 0.9563 0.0440
-4.500 -0.2442 0.01551 0.00764 -0.0608 0.9532 0.0458
-4.250 -0.2103 0.01500 0.00705 -0.0621 0.9509 0.0484
-4.000 -0.1859 0.01461 0.00658 -0.0614 0.9451 0.0510
-3.750 -0.1561 0.01413 0.00607 -0.0618 0.9415 0.0538
-3.500 -0.1238 0.01371 0.00565 -0.0628 0.9388 0.0586
-3.250 -0.0897 0.01333 0.00527 -0.0641 0.9367 0.0662
-3.000 -0.0652 0.01309 0.00505 -0.0633 0.9305 0.0782
-2.750 -0.0342 0.01287 0.00484 -0.0640 0.9267 0.0930
-2.500 -0.0003 0.01266 0.00462 -0.0652 0.9239 0.1062
-2.250 0.0348 0.01245 0.00442 -0.0667 0.9216 0.1196
-2.000 0.0588 0.01229 0.00430 -0.0659 0.9143 0.1311
-1.750 0.0907 0.01206 0.00412 -0.0667 0.9102 0.1443
-1.500 0.1251 0.01181 0.00391 -0.0680 0.9071 0.1584
-1.250 0.1504 0.01162 0.00379 -0.0674 0.8996 0.1766
-1.000 0.1822 0.01130 0.00361 -0.0682 0.8947 0.2080
-0.750 0.2106 0.01091 0.00347 -0.0683 0.8879 0.2751
-0.500 0.2386 0.01012 0.00331 -0.0684 0.8793 0.4617
-0.250 0.2597 0.00922 0.00324 -0.0665 0.8668 0.7003
0.000 0.3345 0.00868 0.00319 -0.0754 0.8553 0.9495
0.250 0.3921 0.00859 0.00303 -0.0815 0.8429 0.9904
0.500 0.4292 0.00856 0.00291 -0.0832 0.8265 1.0000
0.750 0.4567 0.00855 0.00278 -0.0828 0.8056 1.0000
1.000 0.4809 0.00859 0.00270 -0.0817 0.7784 1.0000
1.250 0.5048 0.00865 0.00265 -0.0806 0.7501 1.0000
1.500 0.5285 0.00873 0.00261 -0.0794 0.7178 1.0000
1.750 0.5516 0.00887 0.00258 -0.0782 0.6824 1.0000
2.000 0.5740 0.00906 0.00259 -0.0768 0.6523 1.0000
2.250 0.5959 0.00928 0.00266 -0.0754 0.6253 1.0000
2.500 0.6175 0.00953 0.00276 -0.0739 0.5985 1.0000
2.750 0.6387 0.00978 0.00291 -0.0724 0.5682 1.0000
3.000 0.6595 0.01005 0.00305 -0.0709 0.5308 1.0000
3.250 0.6784 0.01041 0.00320 -0.0689 0.4804 1.0000
3.500 0.6954 0.01089 0.00340 -0.0667 0.4216 1.0000
3.750 0.7124 0.01144 0.00366 -0.0646 0.3633 1.0000
4.000 0.7310 0.01194 0.00394 -0.0629 0.3177 1.0000
4.250 0.7510 0.01240 0.00424 -0.0614 0.2842 1.0000
4.500 0.7717 0.01281 0.00456 -0.0601 0.2605 1.0000
4.750 0.7931 0.01320 0.00489 -0.0589 0.2439 1.0000
5.000 0.8146 0.01358 0.00522 -0.0577 0.2279 1.0000
5.250 0.8362 0.01396 0.00557 -0.0566 0.2144 1.0000
5.500 0.8581 0.01433 0.00594 -0.0555 0.2012 1.0000
5.750 0.8805 0.01467 0.00630 -0.0545 0.1903 1.0000
6.000 0.9026 0.01503 0.00668 -0.0535 0.1787 1.0000
6.250 0.9244 0.01540 0.00707 -0.0524 0.1669 1.0000
6.500 0.9462 0.01579 0.00748 -0.0513 0.1519 1.0000
6.750 0.9676 0.01621 0.00787 -0.0503 0.1304 1.0000
7.250 1.0064 0.01744 0.00886 -0.0476 0.0876 1.0000
7.500 1.0255 0.01810 0.00952 -0.0462 0.0742 1.0000
7.750 1.0448 0.01876 0.01018 -0.0449 0.0513 1.0000
8.000 1.0616 0.01967 0.01096 -0.0432 0.0379 1.0000
8.250 1.0782 0.02059 0.01190 -0.0414 0.0320 1.0000
8.500 1.0948 0.02146 0.01286 -0.0396 0.0278 1.0000
8.750 1.1079 0.02259 0.01407 -0.0374 0.0247 1.0000
9.000 1.1238 0.02342 0.01504 -0.0355 0.0230 1.0000
9.250 1.1382 0.02434 0.01611 -0.0335 0.0210 1.0000
9.500 1.1509 0.02527 0.01714 -0.0312 0.0194 1.0000
9.750 1.1583 0.02655 0.01849 -0.0283 0.0178 1.0000
10.000 1.1668 0.02777 0.01982 -0.0256 0.0171 1.0000
10.250 1.1765 0.02895 0.02115 -0.0231 0.0163 1.0000
10.500 1.1851 0.03025 0.02259 -0.0206 0.0156 1.0000
10.750 1.1931 0.03160 0.02407 -0.0182 0.0149 1.0000
11.000 1.2007 0.03296 0.02556 -0.0160 0.0142 1.0000
11.250 1.2076 0.03433 0.02704 -0.0139 0.0136 1.0000
11.500 1.2122 0.03585 0.02864 -0.0118 0.0129 1.0000
11.750 1.2141 0.03784 0.03079 -0.0096 0.0125 1.0000
12.000 1.2141 0.04020 0.03330 -0.0075 0.0121 1.0000
12.250 1.2159 0.04238 0.03569 -0.0057 0.0119 1.0000
12.500 1.2158 0.04481 0.03832 -0.0042 0.0117 1.0000
12.750 1.2135 0.04752 0.04126 -0.0028 0.0114 1.0000
13.000 1.2097 0.05051 0.04445 -0.0018 0.0113 1.0000
13.250 1.2034 0.05385 0.04802 -0.0013 0.0111 1.0000
13.500 1.1948 0.05768 0.05207 -0.0012 0.0109 1.0000
13.750 1.1851 0.06188 0.05647 -0.0018 0.0109 1.0000
14.000 1.1731 0.06667 0.06149 -0.0031 0.0108 1.0000
14.250 1.1590 0.07212 0.06715 -0.0052 0.0107 1.0000
14.500 1.1442 0.07808 0.07332 -0.0080 0.0107 1.0000
14.750 1.1268 0.08502 0.08046 -0.0117 0.0107 1.0000
15.000 1.1084 0.09275 0.08839 -0.0164 0.0106 1.0000
15.250 1.0863 0.10194 0.09777 -0.0220 0.0107 1.0000
15.500 1.0647 0.11163 0.10764 -0.0280 0.0109 1.0000
15.750 1.0401 0.12261 0.11878 -0.0348 0.0111 1.0000
16.000 1.0133 0.13478 0.13108 -0.0421 0.0113 1.0000
16.250 0.9810 0.14927 0.14567 -0.0504 0.0116 1.0000
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