GOE 595 AIRFOIL (goe595-il) Xfoil prediction polar at RE=200,000 Ncrit=9
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Airfoil: GOE 595 AIRFOIL (goe595-il) Reynolds number: 200,000 Max Cl/Cd: 74.61 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe595-il-200000.txt Download as CSV file: xf-goe595-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 595 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3195 0.08874 0.08548 -0.0358 1.0000 0.0451
-9.000 -0.3239 0.08564 0.08242 -0.0355 1.0000 0.0463
-8.750 -0.3314 0.08249 0.07932 -0.0352 1.0000 0.0480
-8.500 -0.3436 0.07934 0.07623 -0.0352 1.0000 0.0494
-8.250 -0.3602 0.07627 0.07323 -0.0351 1.0000 0.0501
-8.000 -0.3811 0.07350 0.07053 -0.0348 1.0000 0.0505
-7.750 -0.4531 0.08101 0.07787 -0.0313 1.0000 0.0453
-7.500 -0.4698 0.07856 0.07548 -0.0302 1.0000 0.0452
-7.250 -0.4817 0.07556 0.07251 -0.0299 1.0000 0.0461
-7.000 -0.4915 0.07214 0.06909 -0.0302 1.0000 0.0475
-6.750 -0.4981 0.06839 0.06530 -0.0309 1.0000 0.0486
-6.250 -0.5096 0.05787 0.05434 -0.0335 1.0000 0.0523
-6.000 -0.4996 0.05730 0.05396 -0.0309 0.9996 0.0550
-5.750 -0.4694 0.05084 0.04686 -0.0383 0.9943 0.0644
-5.500 -0.4468 0.04826 0.04445 -0.0392 0.9908 0.0672
-5.250 -0.4172 0.04402 0.03958 -0.0431 0.9855 0.0779
-5.000 -0.3908 0.04113 0.03682 -0.0447 0.9824 0.0813
-4.750 -0.3647 0.03794 0.03317 -0.0463 0.9763 0.0929
-4.500 -0.3319 0.02768 0.02160 -0.0453 0.9736 0.0469
-4.250 -0.2947 0.02442 0.01772 -0.0468 0.9709 0.0453
-4.000 -0.2659 0.02181 0.01453 -0.0462 0.9648 0.0431
-3.750 -0.2277 0.02011 0.01255 -0.0478 0.9606 0.0437
-3.500 -0.1852 0.01900 0.01131 -0.0506 0.9575 0.0471
-3.000 -0.1162 0.01685 0.00901 -0.0527 0.9456 0.0534
-2.750 -0.0821 0.01631 0.00848 -0.0539 0.9402 0.0608
-2.500 -0.0515 0.01567 0.00785 -0.0544 0.9345 0.0678
-2.250 -0.0131 0.01504 0.00724 -0.0564 0.9311 0.0894
-2.000 0.0117 0.01386 0.00690 -0.0563 0.9249 0.2695
-1.750 0.0360 0.01270 0.00678 -0.0559 0.9193 0.5399
-1.500 0.0824 0.01171 0.00696 -0.0583 0.9189 0.8861
-1.250 0.2038 0.01162 0.00672 -0.0774 0.9297 0.9897
-1.000 0.2488 0.01155 0.00651 -0.0811 0.9247 1.0000
-0.750 0.2881 0.01138 0.00623 -0.0835 0.9196 1.0000
-0.500 0.3331 0.01111 0.00587 -0.0868 0.9162 1.0000
-0.250 0.3627 0.01093 0.00562 -0.0870 0.9065 1.0000
0.000 0.4073 0.01055 0.00517 -0.0900 0.9012 1.0000
0.250 0.4340 0.01037 0.00494 -0.0895 0.8896 1.0000
0.500 0.4622 0.01019 0.00471 -0.0892 0.8784 1.0000
0.750 0.4913 0.01000 0.00448 -0.0891 0.8672 1.0000
1.000 0.5197 0.00984 0.00427 -0.0888 0.8556 1.0000
1.250 0.5445 0.00971 0.00412 -0.0878 0.8411 1.0000
1.500 0.5675 0.00961 0.00398 -0.0864 0.8240 1.0000
1.750 0.5914 0.00953 0.00385 -0.0851 0.8060 1.0000
2.000 0.6167 0.00947 0.00373 -0.0843 0.7880 1.0000
2.250 0.6405 0.00946 0.00370 -0.0831 0.7678 1.0000
2.500 0.6647 0.00949 0.00365 -0.0820 0.7457 1.0000
2.750 0.6876 0.00957 0.00365 -0.0807 0.7205 1.0000
3.000 0.7100 0.00971 0.00369 -0.0793 0.6932 1.0000
3.250 0.7315 0.00989 0.00378 -0.0778 0.6642 1.0000
3.500 0.7525 0.01011 0.00390 -0.0762 0.6352 1.0000
3.750 0.7730 0.01036 0.00404 -0.0746 0.6068 1.0000
4.000 0.7926 0.01065 0.00420 -0.0727 0.5766 1.0000
4.250 0.8113 0.01099 0.00440 -0.0708 0.5458 1.0000
4.500 0.8296 0.01135 0.00462 -0.0688 0.5155 1.0000
4.750 0.8484 0.01171 0.00488 -0.0670 0.4882 1.0000
5.000 0.8681 0.01207 0.00517 -0.0653 0.4653 1.0000
5.250 0.8875 0.01244 0.00548 -0.0637 0.4423 1.0000
5.500 0.9062 0.01280 0.00582 -0.0619 0.4159 1.0000
5.750 0.9242 0.01317 0.00614 -0.0600 0.3863 1.0000
6.000 0.9422 0.01353 0.00645 -0.0582 0.3551 1.0000
6.250 0.9592 0.01394 0.00678 -0.0561 0.3152 1.0000
6.500 0.9721 0.01463 0.00721 -0.0535 0.2620 1.0000
6.750 0.9815 0.01570 0.00786 -0.0505 0.1893 1.0000
7.000 0.9832 0.01763 0.00896 -0.0464 0.0764 1.0000
7.250 0.9955 0.01873 0.01005 -0.0438 0.0627 1.0000
7.500 1.0049 0.01994 0.01129 -0.0407 0.0552 1.0000
7.750 1.0166 0.02093 0.01239 -0.0380 0.0502 1.0000
8.000 1.0223 0.02223 0.01371 -0.0344 0.0464 1.0000
8.250 1.0290 0.02376 0.01529 -0.0311 0.0431 1.0000
8.500 1.0425 0.02483 0.01645 -0.0288 0.0400 1.0000
8.750 1.0572 0.02618 0.01785 -0.0269 0.0376 1.0000
9.000 1.0771 0.02814 0.01977 -0.0261 0.0350 1.0000
9.250 1.1058 0.03078 0.02254 -0.0268 0.0324 1.0000
9.500 1.1253 0.03223 0.02418 -0.0255 0.0308 1.0000
9.750 1.1478 0.03435 0.02654 -0.0248 0.0296 1.0000
10.000 1.1680 0.03679 0.02922 -0.0238 0.0290 1.0000
10.250 1.1835 0.03933 0.03204 -0.0222 0.0284 1.0000
10.500 1.1954 0.04154 0.03444 -0.0204 0.0274 1.0000
10.750 1.2071 0.04368 0.03663 -0.0190 0.0259 1.0000
11.000 1.2145 0.04726 0.04045 -0.0172 0.0254 1.0000
11.250 1.2145 0.04985 0.04333 -0.0138 0.0255 1.0000
11.500 1.2054 0.05231 0.04616 -0.0091 0.0262 1.0000
11.750 1.1734 0.05594 0.05037 -0.0028 0.0279 1.0000
12.000 1.1532 0.05980 0.05456 0.0006 0.0289 1.0000
12.250 1.1322 0.06425 0.05927 0.0029 0.0302 1.0000
12.500 1.1120 0.06859 0.06383 0.0042 0.0308 1.0000
12.750 1.0917 0.07359 0.06901 0.0044 0.0316 1.0000
13.000 0.9729 0.07337 0.06923 0.0062 0.0300 1.0000
13.250 0.9455 0.07983 0.07586 0.0041 0.0303 1.0000
13.500 0.9173 0.08670 0.08290 0.0011 0.0306 1.0000
13.750 0.8878 0.09400 0.09036 -0.0030 0.0309 1.0000
14.000 0.8562 0.10174 0.09826 -0.0079 0.0312 1.0000
14.250 0.8225 0.10929 0.10593 -0.0131 0.0322 1.0000
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Polar data table (+)
Polar graphs
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