GOE 767 AIRFOIL (goe767-il) Xfoil prediction polar at RE=50,000 Ncrit=5
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Airfoil: GOE 767 AIRFOIL (goe767-il) Reynolds number: 50,000 Max Cl/Cd: 23.94 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe767-il-50000-n5.txt Download as CSV file: xf-goe767-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 767 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.6192 0.09237 0.08433 -0.0031 1.0000 0.0877
-10.000 -0.6460 0.08224 0.07423 -0.0105 1.0000 0.0880
-9.750 -0.6903 0.07150 0.06344 -0.0185 1.0000 0.0877
-9.500 -0.7282 0.06455 0.05631 -0.0204 1.0000 0.0879
-9.250 -0.7565 0.05809 0.04949 -0.0210 1.0000 0.0887
-9.000 -0.7579 0.05454 0.04574 -0.0205 1.0000 0.0903
-8.750 -0.7446 0.05283 0.04402 -0.0196 1.0000 0.0925
-8.500 -0.7395 0.04978 0.04074 -0.0188 1.0000 0.0948
-8.250 -0.7368 0.04599 0.03652 -0.0178 1.0000 0.0976
-8.000 -0.7295 0.04255 0.03263 -0.0166 1.0000 0.1008
-7.750 -0.7103 0.04131 0.03141 -0.0157 1.0000 0.1041
-7.500 -0.6951 0.03920 0.02902 -0.0146 1.0000 0.1090
-7.250 -0.6785 0.03714 0.02671 -0.0135 1.0000 0.1140
-7.000 -0.6584 0.03593 0.02547 -0.0125 1.0000 0.1194
-6.750 -0.6401 0.03410 0.02329 -0.0114 1.0000 0.1263
-6.500 -0.6185 0.03321 0.02248 -0.0105 1.0000 0.1326
-6.250 -0.5983 0.03188 0.02090 -0.0094 1.0000 0.1408
-6.000 -0.5763 0.03107 0.02016 -0.0085 1.0000 0.1485
-5.750 -0.5548 0.03002 0.01895 -0.0075 1.0000 0.1571
-5.500 -0.5323 0.02915 0.01808 -0.0065 1.0000 0.1653
-5.250 -0.5096 0.02813 0.01697 -0.0056 1.0000 0.1731
-5.000 -0.4862 0.02723 0.01608 -0.0047 1.0000 0.1802
-4.750 -0.4624 0.02632 0.01501 -0.0038 1.0000 0.1877
-4.500 -0.4384 0.02545 0.01427 -0.0030 1.0000 0.1941
-4.250 -0.4142 0.02468 0.01343 -0.0021 1.0000 0.2019
-4.000 -0.3900 0.02390 0.01276 -0.0013 1.0000 0.2091
-3.750 -0.3658 0.02322 0.01214 -0.0004 1.0000 0.2170
-3.500 -0.3417 0.02258 0.01158 0.0004 1.0000 0.2259
-3.250 -0.3182 0.02200 0.01113 0.0012 1.0000 0.2362
-3.000 -0.2927 0.02143 0.01074 0.0014 0.9951 0.2475
-2.750 -0.2366 0.02073 0.01022 -0.0041 0.9297 0.2718
-2.500 -0.1832 0.02004 0.00971 -0.0085 0.8462 0.3072
-2.250 -0.1448 0.01948 0.00926 -0.0096 0.7637 0.3592
-2.000 -0.1203 0.01904 0.00898 -0.0080 0.7060 0.4327
-1.750 -0.1002 0.01869 0.00890 -0.0054 0.6651 0.5320
-1.500 -0.0783 0.01853 0.00901 -0.0026 0.6324 0.6450
-1.250 -0.0498 0.01861 0.00917 -0.0009 0.6037 0.7438
-1.000 -0.0111 0.01886 0.00937 -0.0010 0.5762 0.8254
-0.750 0.0355 0.01923 0.00954 -0.0030 0.5506 0.8882
-0.500 0.0876 0.01961 0.00964 -0.0064 0.5271 0.9385
-0.250 0.1573 0.01987 0.00961 -0.0137 0.5020 0.9831
0.000 0.2009 0.01997 0.00947 -0.0170 0.4836 1.0000
0.250 0.2215 0.02011 0.00946 -0.0159 0.4703 1.0000
0.500 0.2429 0.02029 0.00953 -0.0149 0.4573 1.0000
0.750 0.2644 0.02051 0.00960 -0.0138 0.4463 1.0000
1.000 0.2860 0.02076 0.00975 -0.0128 0.4345 1.0000
1.250 0.3083 0.02105 0.00994 -0.0118 0.4235 1.0000
1.500 0.3309 0.02136 0.01012 -0.0108 0.4138 1.0000
1.750 0.3537 0.02173 0.01045 -0.0099 0.4041 1.0000
2.000 0.3764 0.02210 0.01072 -0.0089 0.3953 1.0000
2.250 0.3992 0.02253 0.01111 -0.0080 0.3864 1.0000
2.500 0.4220 0.02295 0.01147 -0.0070 0.3779 1.0000
2.750 0.4449 0.02342 0.01187 -0.0061 0.3703 1.0000
3.000 0.4676 0.02394 0.01240 -0.0052 0.3618 1.0000
3.250 0.4907 0.02436 0.01268 -0.0041 0.3551 1.0000
3.500 0.5129 0.02497 0.01340 -0.0032 0.3456 1.0000
3.750 0.5357 0.02539 0.01372 -0.0021 0.3383 1.0000
4.000 0.5577 0.02600 0.01440 -0.0012 0.3290 1.0000
4.250 0.5804 0.02642 0.01475 -0.0001 0.3211 1.0000
4.500 0.6023 0.02704 0.01544 0.0008 0.3124 1.0000
4.750 0.6248 0.02754 0.01591 0.0018 0.3043 1.0000
5.000 0.6470 0.02817 0.01656 0.0028 0.2967 1.0000
5.250 0.6687 0.02884 0.01731 0.0037 0.2884 1.0000
5.500 0.6917 0.02933 0.01771 0.0047 0.2819 1.0000
5.750 0.7119 0.03029 0.01887 0.0056 0.2731 1.0000
6.000 0.7353 0.03072 0.01920 0.0066 0.2670 1.0000
6.250 0.7541 0.03188 0.02060 0.0075 0.2584 1.0000
6.500 0.7764 0.03250 0.02119 0.0084 0.2519 1.0000
6.750 0.7956 0.03361 0.02243 0.0093 0.2447 1.0000
7.000 0.8152 0.03460 0.02353 0.0103 0.2375 1.0000
7.250 0.8388 0.03505 0.02385 0.0112 0.2325 1.0000
7.500 0.8511 0.03711 0.02628 0.0121 0.2247 1.0000
7.750 0.8717 0.03792 0.02712 0.0130 0.2192 1.0000
8.000 0.8890 0.03923 0.02851 0.0140 0.2138 1.0000
8.250 0.8982 0.04147 0.03104 0.0149 0.2072 1.0000
8.500 0.9177 0.04240 0.03197 0.0158 0.2027 1.0000
8.750 0.9315 0.04404 0.03372 0.0167 0.1982 1.0000
9.000 0.9253 0.04779 0.03786 0.0175 0.1925 1.0000
9.250 0.9366 0.04940 0.03954 0.0184 0.1881 1.0000
9.500 0.9655 0.04923 0.03925 0.0195 0.1848 1.0000
9.750 0.8960 0.05902 0.04956 0.0185 0.1807 1.0000
10.000 0.7742 0.07994 0.07056 0.0053 0.1745 1.0000
10.250 0.7942 0.08003 0.07066 0.0068 0.1718 1.0000
10.750 0.7286 0.09933 0.08994 -0.0042 0.1621 1.0000
11.000 0.7394 0.10133 0.09196 -0.0041 0.1588 1.0000
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Polar data table (+)
Polar graphs
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