GOE 459 AIRFOIL (goe459-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: GOE 459 AIRFOIL (goe459-il) Reynolds number: 50,000 Max Cl/Cd: 28.6 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe459-il-50000.txt Download as CSV file: xf-goe459-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 459 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
------ -------- --------- --------- -------- -------- --------
-10.500 -0.6495 0.09094 0.08311 -0.0246 1.0000 0.2097
-10.250 -0.6543 0.08497 0.07715 -0.0262 1.0000 0.2063
-10.000 -0.8429 0.06711 0.05930 -0.0290 1.0000 0.1800
-9.750 -0.8615 0.06286 0.05487 -0.0259 1.0000 0.1796
-9.500 -0.8797 0.05886 0.05060 -0.0223 1.0000 0.1796
-9.250 -0.8962 0.05520 0.04658 -0.0183 1.0000 0.1802
-9.000 -0.8707 0.05300 0.04456 -0.0178 1.0000 0.1864
-8.750 -0.8746 0.04985 0.04112 -0.0145 1.0000 0.1883
-8.500 -0.8761 0.04678 0.03768 -0.0112 1.0000 0.1901
-8.250 -0.8756 0.04397 0.03445 -0.0078 1.0000 0.1928
-8.000 -0.8700 0.04150 0.03165 -0.0048 1.0000 0.1972
-7.750 -0.8524 0.03964 0.02977 -0.0032 1.0000 0.2034
-7.500 -0.8426 0.03746 0.02720 -0.0005 1.0000 0.2087
-7.250 -0.8248 0.03544 0.02506 0.0011 1.0000 0.2147
-7.000 -0.8087 0.03376 0.02316 0.0031 1.0000 0.2237
-6.750 -0.7897 0.03224 0.02167 0.0047 1.0000 0.2356
-6.500 -0.7714 0.03071 0.02012 0.0065 1.0000 0.2506
-6.250 -0.7535 0.02915 0.01851 0.0084 1.0000 0.2708
-6.000 -0.7355 0.02761 0.01716 0.0104 1.0000 0.3012
-5.750 -0.7208 0.02621 0.01613 0.0131 1.0000 0.3432
-5.500 -0.7081 0.02517 0.01544 0.0163 1.0000 0.3888
-5.250 -0.6960 0.02446 0.01500 0.0197 1.0000 0.4342
-5.000 -0.6840 0.02391 0.01468 0.0232 1.0000 0.4804
-4.750 -0.6716 0.02356 0.01455 0.0270 1.0000 0.5285
-4.500 -0.6582 0.02355 0.01484 0.0312 1.0000 0.5790
-4.250 -0.6424 0.02381 0.01533 0.0355 1.0000 0.6297
-4.000 -0.6224 0.02415 0.01578 0.0392 1.0000 0.6775
-3.750 -0.5981 0.02452 0.01618 0.0421 1.0000 0.7216
-3.500 -0.5619 0.02525 0.01685 0.0433 1.0000 0.7619
-3.250 -0.5240 0.02595 0.01740 0.0437 1.0000 0.7992
-3.000 -0.4510 0.02736 0.01856 0.0386 1.0000 0.8331
-2.750 -0.3467 0.02883 0.01966 0.0277 1.0000 0.8657
-2.500 -0.2662 0.02930 0.01986 0.0192 1.0000 0.8961
-2.250 -0.1873 0.02914 0.01949 0.0098 1.0000 0.9211
-2.000 -0.1261 0.02866 0.01887 0.0027 1.0000 0.9451
-1.750 -0.0627 0.02797 0.01807 -0.0053 1.0000 0.9683
-1.500 0.0087 0.02692 0.01692 -0.0153 1.0000 0.9906
-1.250 0.0424 0.02618 0.01615 -0.0189 1.0000 1.0000
-1.000 0.0399 0.02590 0.01592 -0.0159 1.0000 1.0000
-0.750 0.0337 0.02572 0.01578 -0.0124 1.0000 1.0000
-0.500 0.0243 0.02562 0.01571 -0.0085 1.0000 1.0000
-0.250 0.0127 0.02558 0.01568 -0.0043 1.0000 1.0000
0.000 0.0000 0.02557 0.01567 0.0000 1.0000 1.0000
0.250 -0.0127 0.02558 0.01567 0.0043 1.0000 1.0000
0.500 -0.0243 0.02562 0.01570 0.0085 1.0000 1.0000
0.750 -0.0337 0.02572 0.01577 0.0124 1.0000 1.0000
1.000 -0.0399 0.02590 0.01592 0.0159 1.0000 1.0000
1.250 -0.0424 0.02617 0.01615 0.0189 1.0000 1.0000
1.500 -0.0087 0.02692 0.01691 0.0153 0.9905 1.0000
1.750 0.0628 0.02797 0.01806 0.0053 0.9683 1.0000
2.000 0.1260 0.02865 0.01886 -0.0027 0.9452 1.0000
2.250 0.1873 0.02913 0.01948 -0.0098 0.9211 1.0000
2.500 0.2664 0.02929 0.01985 -0.0192 0.8961 1.0000
2.750 0.3461 0.02883 0.01966 -0.0276 0.8658 1.0000
3.000 0.4517 0.02733 0.01853 -0.0387 0.8330 1.0000
3.250 0.5238 0.02595 0.01740 -0.0437 0.7993 1.0000
3.500 0.5618 0.02524 0.01684 -0.0433 0.7619 1.0000
3.750 0.5981 0.02451 0.01617 -0.0421 0.7216 1.0000
4.000 0.6223 0.02415 0.01579 -0.0392 0.6777 1.0000
4.250 0.6423 0.02380 0.01532 -0.0355 0.6297 1.0000
4.500 0.6581 0.02355 0.01484 -0.0312 0.5791 1.0000
4.750 0.6715 0.02355 0.01455 -0.0270 0.5285 1.0000
5.000 0.6839 0.02391 0.01468 -0.0232 0.4805 1.0000
5.250 0.6960 0.02446 0.01500 -0.0197 0.4344 1.0000
5.500 0.7081 0.02517 0.01544 -0.0163 0.3890 1.0000
5.750 0.7206 0.02620 0.01612 -0.0130 0.3429 1.0000
6.000 0.7356 0.02760 0.01715 -0.0104 0.3014 1.0000
6.250 0.7536 0.02915 0.01850 -0.0084 0.2709 1.0000
6.500 0.7713 0.03072 0.02012 -0.0065 0.2505 1.0000
6.750 0.7899 0.03225 0.02167 -0.0047 0.2359 1.0000
7.000 0.8087 0.03376 0.02316 -0.0031 0.2237 1.0000
7.250 0.8248 0.03544 0.02507 -0.0011 0.2149 1.0000
7.500 0.8426 0.03746 0.02720 0.0005 0.2087 1.0000
7.750 0.8524 0.03965 0.02978 0.0032 0.2035 1.0000
8.000 0.8697 0.04151 0.03168 0.0048 0.1974 1.0000
8.250 0.8759 0.04397 0.03445 0.0077 0.1929 1.0000
8.500 0.8761 0.04678 0.03768 0.0112 0.1901 1.0000
8.750 0.8747 0.04985 0.04111 0.0145 0.1883 1.0000
9.000 0.8708 0.05301 0.04456 0.0178 0.1864 1.0000
9.250 0.8962 0.05517 0.04655 0.0182 0.1801 1.0000
9.500 0.8793 0.05886 0.05061 0.0224 0.1795 1.0000
9.750 0.8610 0.06287 0.05488 0.0260 0.1795 1.0000
10.000 0.8420 0.06713 0.05932 0.0290 0.1799 1.0000
10.250 0.8218 0.07162 0.06392 0.0315 0.1804 1.0000
10.500 0.4938 0.11489 0.10661 -0.0012 0.3744 1.0000
10.750 0.5229 0.11987 0.11169 -0.0013 0.3680 1.0000
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Polar data table (+)
Polar graphs
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