GOE 388 AIRFOIL (goe388-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 388 AIRFOIL (goe388-il) Reynolds number: 50,000 Max Cl/Cd: 41.36 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe388-il-50000.txt Download as CSV file: xf-goe388-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 388 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3470 0.10288 0.09691 -0.0296 1.0000 0.1969
-7.500 -0.3750 0.10386 0.09809 -0.0267 1.0000 0.1989
-7.250 -0.4030 0.10522 0.09961 -0.0269 1.0000 0.2000
-7.000 -0.3757 0.09814 0.09251 -0.0216 1.0000 0.2080
-6.750 -0.3909 0.09749 0.09197 -0.0199 1.0000 0.2134
-6.500 -0.4179 0.09907 0.09367 -0.0234 1.0000 0.2167
-6.250 -0.4036 0.09342 0.08806 -0.0165 1.0000 0.2227
-6.000 -0.4119 0.09227 0.08697 -0.0166 1.0000 0.2311
-5.750 -0.4150 0.08971 0.08447 -0.0162 1.0000 0.2365
-5.500 -0.4138 0.08748 0.08227 -0.0146 1.0000 0.2463
-5.250 -0.4139 0.08490 0.07974 -0.0147 1.0000 0.2544
-5.000 -0.3807 0.08075 0.07551 -0.0197 0.9910 0.2776
-4.750 -0.3449 0.07642 0.07111 -0.0263 0.9804 0.3048
-4.500 -0.3157 0.07314 0.06777 -0.0324 0.9703 0.3347
-4.250 -0.2923 0.06939 0.06402 -0.0319 0.9626 0.3577
-4.000 -0.2675 0.06614 0.06075 -0.0335 0.9545 0.3920
-3.750 -0.2415 0.06276 0.05735 -0.0340 0.9424 0.4379
-3.500 -0.2107 0.05892 0.05352 -0.0317 0.9272 0.5051
-2.500 0.1115 0.03762 0.02881 -0.1097 0.8924 0.1828
-2.250 0.1542 0.03509 0.02602 -0.1132 0.8884 0.1773
-2.000 0.1836 0.03368 0.02424 -0.1141 0.8805 0.1783
-1.750 0.2206 0.03216 0.02233 -0.1161 0.8752 0.1795
-1.500 0.2602 0.03070 0.02051 -0.1183 0.8715 0.1797
-1.250 0.2876 0.02979 0.01929 -0.1181 0.8613 0.1820
-1.000 0.3250 0.02831 0.01765 -0.1190 0.8499 0.1884
-0.750 0.3624 0.02715 0.01634 -0.1199 0.8404 0.2045
-0.500 0.3902 0.02641 0.01562 -0.1195 0.8304 0.2268
-0.250 0.4235 0.02540 0.01477 -0.1201 0.8250 0.2694
0.000 0.4487 0.02252 0.01371 -0.1183 0.8198 1.0000
0.250 0.4742 0.02284 0.01364 -0.1179 0.8123 1.0000
0.500 0.5052 0.02302 0.01355 -0.1186 0.8077 1.0000
0.750 0.5276 0.02347 0.01384 -0.1180 0.7995 1.0000
1.000 0.5561 0.02373 0.01395 -0.1184 0.7938 1.0000
1.250 0.5830 0.02406 0.01417 -0.1185 0.7876 1.0000
1.500 0.6080 0.02441 0.01445 -0.1182 0.7794 1.0000
1.750 0.6367 0.02449 0.01447 -0.1182 0.7710 1.0000
2.000 0.6665 0.02432 0.01423 -0.1179 0.7601 1.0000
2.250 0.6903 0.02454 0.01443 -0.1171 0.7486 1.0000
2.500 0.7228 0.02442 0.01431 -0.1174 0.7420 1.0000
2.750 0.7436 0.02489 0.01482 -0.1165 0.7304 1.0000
3.000 0.7722 0.02506 0.01502 -0.1165 0.7236 1.0000
3.250 0.7948 0.02551 0.01555 -0.1158 0.7134 1.0000
3.500 0.8194 0.02567 0.01580 -0.1150 0.7016 1.0000
3.750 0.8485 0.02572 0.01594 -0.1148 0.6936 1.0000
4.000 0.8722 0.02564 0.01593 -0.1134 0.6776 1.0000
4.250 0.8938 0.02538 0.01572 -0.1114 0.6564 1.0000
4.500 0.9141 0.02535 0.01582 -0.1095 0.6366 1.0000
4.750 0.9359 0.02519 0.01574 -0.1077 0.6160 1.0000
5.000 0.9593 0.02501 0.01561 -0.1061 0.5964 1.0000
5.250 0.9797 0.02503 0.01568 -0.1042 0.5727 1.0000
5.500 1.0028 0.02512 0.01587 -0.1029 0.5553 1.0000
5.750 1.0256 0.02528 0.01607 -0.1015 0.5374 1.0000
6.000 1.0477 0.02554 0.01642 -0.1001 0.5205 1.0000
6.250 1.0688 0.02592 0.01692 -0.0987 0.5044 1.0000
6.500 1.0887 0.02632 0.01746 -0.0970 0.4844 1.0000
6.750 1.1055 0.02688 0.01818 -0.0950 0.4592 1.0000
7.000 1.1209 0.02749 0.01888 -0.0927 0.4239 1.0000
7.250 1.1313 0.02830 0.01935 -0.0895 0.3574 1.0000
7.500 1.1348 0.02977 0.02036 -0.0858 0.2890 1.0000
7.750 1.1355 0.03162 0.02185 -0.0823 0.2320 1.0000
8.000 1.1377 0.03359 0.02356 -0.0793 0.2096 1.0000
8.250 1.1428 0.03545 0.02535 -0.0767 0.2001 1.0000
8.500 1.1505 0.03721 0.02707 -0.0744 0.1944 1.0000
8.750 1.1609 0.03888 0.02872 -0.0721 0.1880 1.0000
9.000 1.1776 0.04038 0.03010 -0.0700 0.1824 1.0000
9.250 1.2170 0.04138 0.03114 -0.0694 0.1744 1.0000
9.500 1.3093 0.04459 0.03392 -0.0756 0.1519 1.0000
9.750 1.3601 0.04897 0.03823 -0.0784 0.1392 1.0000
10.000 1.3961 0.05426 0.04343 -0.0801 0.1271 1.0000
10.250 1.3974 0.05650 0.04627 -0.0766 0.1219 1.0000
10.500 1.4231 0.06203 0.05178 -0.0773 0.1129 1.0000
10.750 1.4187 0.06440 0.05473 -0.0734 0.1097 1.0000
11.000 1.4384 0.07003 0.06028 -0.0738 0.1013 1.0000
11.250 1.4233 0.07213 0.06300 -0.0691 0.0993 1.0000
11.500 1.4119 0.07498 0.06626 -0.0654 0.0962 1.0000
11.750 1.4159 0.07946 0.07079 -0.0643 0.0899 1.0000
12.000 1.3958 0.08226 0.07399 -0.0602 0.0887 1.0000
12.250 1.3749 0.08521 0.07725 -0.0564 0.0878 1.0000
12.500 1.3532 0.08848 0.08078 -0.0532 0.0871 1.0000
12.750 1.3303 0.09221 0.08474 -0.0509 0.0870 1.0000
13.000 1.3063 0.09645 0.08922 -0.0494 0.0874 1.0000
13.250 1.2818 0.10126 0.09424 -0.0490 0.0881 1.0000
13.500 1.2584 0.10661 0.09978 -0.0494 0.0889 1.0000
13.750 1.2362 0.11246 0.10578 -0.0507 0.0897 1.0000
14.000 1.2160 0.11869 0.11214 -0.0526 0.0904 1.0000
14.250 1.1978 0.12527 0.11881 -0.0552 0.0909 1.0000
14.500 0.7623 0.17597 0.16973 -0.0835 0.3382 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 388 AIRFOIL (goe388-il)