GOE 575 AIRFOIL (goe575-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 575 AIRFOIL (goe575-il) Reynolds number: 50,000 Max Cl/Cd: 32.46 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe575-il-50000-n5.txt Download as CSV file: xf-goe575-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 575 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4582 0.10228 0.09483 -0.0310 1.0000 0.0903
-8.750 -0.4718 0.09785 0.09046 -0.0319 1.0000 0.0894
-8.500 -0.4906 0.09333 0.08601 -0.0326 1.0000 0.0884
-8.250 -0.5173 0.08877 0.08150 -0.0317 1.0000 0.0874
-8.000 -0.5462 0.08510 0.07785 -0.0288 1.0000 0.0869
-7.750 -0.5714 0.08132 0.07406 -0.0254 1.0000 0.0867
-7.500 -0.5956 0.07702 0.06969 -0.0220 1.0000 0.0862
-7.250 -0.6064 0.07417 0.06679 -0.0186 1.0000 0.0869
-7.000 -0.6198 0.07084 0.06337 -0.0149 1.0000 0.0872
-6.750 -0.6322 0.06749 0.05988 -0.0109 1.0000 0.0875
-6.500 -0.6420 0.06437 0.05661 -0.0067 1.0000 0.0880
-6.250 -0.6517 0.06115 0.05318 -0.0022 1.0000 0.0881
-6.000 -0.6591 0.05820 0.05000 0.0023 1.0000 0.0882
-5.750 -0.6647 0.05543 0.04698 0.0069 1.0000 0.0886
-5.500 -0.6683 0.05294 0.04423 0.0115 1.0000 0.0891
-5.250 -0.6699 0.05073 0.04175 0.0160 1.0000 0.0900
-5.000 -0.6705 0.04867 0.03938 0.0205 1.0000 0.0913
-4.750 -0.6570 0.04634 0.03652 0.0224 0.9949 0.0933
-4.500 -0.6344 0.04406 0.03363 0.0229 0.9869 0.0947
-4.250 -0.6069 0.04265 0.03200 0.0225 0.9789 0.0959
-4.000 -0.5785 0.04163 0.03082 0.0219 0.9707 0.0979
-3.750 -0.5497 0.04061 0.02958 0.0214 0.9622 0.1010
-3.500 -0.5192 0.03930 0.02791 0.0207 0.9539 0.1047
-3.250 -0.4824 0.03795 0.02618 0.0190 0.9466 0.1079
-3.000 -0.4492 0.03709 0.02528 0.0176 0.9376 0.1110
-2.750 -0.4133 0.03627 0.02434 0.0158 0.9286 0.1159
-2.500 -0.3691 0.03540 0.02325 0.0125 0.9215 0.1234
-2.250 -0.3274 0.03464 0.02247 0.0095 0.9124 0.1314
-2.000 -0.2712 0.03390 0.02169 0.0036 0.9051 0.1449
-1.750 -0.2103 0.03321 0.02091 -0.0032 0.8977 0.1703
-1.500 -0.1570 0.03257 0.02033 -0.0086 0.8889 0.1971
-1.250 -0.1056 0.03197 0.01980 -0.0135 0.8796 0.2299
-1.000 -0.0624 0.03135 0.01932 -0.0168 0.8671 0.2641
-0.750 -0.0049 0.03053 0.01877 -0.0228 0.8595 0.3089
-0.500 0.2051 0.03021 0.02088 -0.0594 0.8659 1.0000
-0.250 0.2380 0.02987 0.02027 -0.0600 0.8498 1.0000
0.000 0.2706 0.02952 0.01970 -0.0606 0.8341 1.0000
0.250 0.3041 0.02913 0.01913 -0.0612 0.8194 1.0000
0.500 0.3390 0.02865 0.01848 -0.0620 0.8051 1.0000
0.750 0.3739 0.02819 0.01788 -0.0628 0.7905 1.0000
1.000 0.4095 0.02772 0.01727 -0.0637 0.7742 1.0000
1.250 0.4473 0.02726 0.01667 -0.0649 0.7580 1.0000
1.500 0.4882 0.02678 0.01604 -0.0668 0.7412 1.0000
1.750 0.5316 0.02632 0.01542 -0.0691 0.7233 1.0000
2.000 0.5750 0.02599 0.01492 -0.0717 0.7046 1.0000
2.250 0.6183 0.02578 0.01454 -0.0742 0.6841 1.0000
2.500 0.6576 0.02576 0.01434 -0.0761 0.6631 1.0000
2.750 0.6894 0.02593 0.01440 -0.0767 0.6418 1.0000
3.000 0.7195 0.02615 0.01450 -0.0770 0.6212 1.0000
3.250 0.7484 0.02641 0.01467 -0.0771 0.6015 1.0000
3.500 0.7762 0.02671 0.01487 -0.0770 0.5827 1.0000
3.750 0.8021 0.02703 0.01509 -0.0765 0.5641 1.0000
4.000 0.8267 0.02737 0.01535 -0.0757 0.5458 1.0000
4.250 0.8486 0.02773 0.01561 -0.0744 0.5276 1.0000
4.500 0.8712 0.02811 0.01592 -0.0733 0.5107 1.0000
4.750 0.8940 0.02849 0.01622 -0.0723 0.4951 1.0000
5.000 0.9172 0.02889 0.01653 -0.0713 0.4799 1.0000
5.250 0.9360 0.02936 0.01703 -0.0696 0.4661 1.0000
5.500 0.9561 0.02985 0.01753 -0.0682 0.4531 1.0000
5.750 0.9783 0.03033 0.01796 -0.0671 0.4405 1.0000
6.000 0.9985 0.03083 0.01845 -0.0657 0.4285 1.0000
6.250 1.0151 0.03140 0.01909 -0.0637 0.4167 1.0000
6.500 1.0366 0.03193 0.01958 -0.0625 0.4058 1.0000
6.750 1.0522 0.03252 0.02022 -0.0603 0.3944 1.0000
7.000 1.0689 0.03317 0.02093 -0.0583 0.3839 1.0000
7.250 1.0893 0.03373 0.02147 -0.0570 0.3738 1.0000
7.500 1.1002 0.03449 0.02234 -0.0541 0.3636 1.0000
7.750 1.1213 0.03508 0.02288 -0.0529 0.3539 1.0000
8.000 1.1290 0.03591 0.02388 -0.0494 0.3441 1.0000
8.250 1.1491 0.03657 0.02451 -0.0482 0.3353 1.0000
8.500 1.1558 0.03749 0.02559 -0.0446 0.3261 1.0000
8.750 1.1763 0.03817 0.02624 -0.0435 0.3179 1.0000
9.000 1.1788 0.03921 0.02749 -0.0393 0.3092 1.0000
9.250 1.1969 0.03983 0.02803 -0.0377 0.3006 1.0000
9.500 1.1932 0.04095 0.02938 -0.0328 0.2916 1.0000
9.750 1.2045 0.04153 0.02989 -0.0301 0.2827 1.0000
10.000 1.1986 0.04268 0.03124 -0.0250 0.2736 1.0000
10.250 1.2047 0.04338 0.03191 -0.0218 0.2649 1.0000
10.500 1.2001 0.04466 0.03337 -0.0173 0.2564 1.0000
10.750 1.2049 0.04551 0.03421 -0.0142 0.2484 1.0000
11.000 1.2007 0.04705 0.03596 -0.0102 0.2408 1.0000
11.250 1.2091 0.04788 0.03672 -0.0078 0.2338 1.0000
11.500 1.1994 0.04992 0.03902 -0.0038 0.2265 1.0000
11.750 1.2043 0.05106 0.04019 -0.0013 0.2200 1.0000
12.000 1.1990 0.05315 0.04245 0.0017 0.2139 1.0000
12.250 1.1948 0.05515 0.04457 0.0043 0.2075 1.0000
12.500 1.1986 0.05655 0.04596 0.0064 0.2014 1.0000
12.750 1.1829 0.05985 0.04951 0.0089 0.1956 1.0000
13.000 1.1847 0.06168 0.05138 0.0106 0.1900 1.0000
13.250 1.1794 0.06436 0.05414 0.0122 0.1848 1.0000
13.500 1.1579 0.06894 0.05897 0.0134 0.1800 1.0000
13.750 1.1570 0.07138 0.06147 0.0144 0.1749 1.0000
14.000 1.1531 0.07435 0.06449 0.0152 0.1702 1.0000
14.250 1.1074 0.08308 0.07352 0.0139 0.1665 1.0000
14.500 1.0621 0.09281 0.08346 0.0113 0.1620 1.0000
14.750 1.1180 0.08692 0.07738 0.0153 0.1569 1.0000
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Polar data table (+)
Polar graphs
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