GOE 746 AIRFOIL (goe746-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 746 AIRFOIL (goe746-il) Reynolds number: 1,000,000 Max Cl/Cd: 118.59 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe746-il-1000000-n5.txt Download as CSV file: xf-goe746-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 746 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4398 0.09439 0.09140 0.0167 0.5127 0.0063
-8.500 -0.4483 0.08913 0.08617 0.0143 0.5113 0.0066
-8.250 -0.4493 0.08601 0.08308 0.0127 0.5095 0.0067
-8.000 -0.4507 0.08344 0.08053 0.0116 0.5074 0.0068
-7.750 -0.4453 0.08071 0.07779 0.0098 0.5050 0.0069
-7.500 -0.4396 0.07714 0.07421 0.0073 0.5027 0.0070
-7.250 -0.4308 0.07372 0.07076 0.0049 0.5001 0.0073
-7.000 -0.4212 0.06980 0.06680 0.0022 0.4976 0.0076
-6.750 -0.4202 0.06069 0.05758 -0.0034 0.4967 0.0089
-6.500 -0.4029 0.05762 0.05446 -0.0051 0.4946 0.0091
-6.250 -0.3840 0.05484 0.05161 -0.0066 0.4926 0.0093
-6.000 -0.3644 0.05189 0.04858 -0.0080 0.4903 0.0096
-5.750 -0.3447 0.04825 0.04483 -0.0094 0.4880 0.0102
-5.500 -0.3297 0.03926 0.03553 -0.0115 0.4865 0.0119
-5.000 -0.2981 0.02444 0.01978 -0.0102 0.4830 0.0152
-4.750 -0.2721 0.02405 0.01932 -0.0100 0.4802 0.0154
-4.500 -0.2458 0.02372 0.01895 -0.0099 0.4780 0.0156
-4.250 -0.2203 0.02290 0.01803 -0.0096 0.4756 0.0160
-4.000 -0.1956 0.02143 0.01638 -0.0091 0.4733 0.0165
-3.750 -0.1749 0.01681 0.01107 -0.0072 0.4712 0.0190
-3.500 -0.1497 0.01543 0.00942 -0.0067 0.4685 0.0195
-3.250 -0.1231 0.01490 0.00880 -0.0064 0.4657 0.0198
-3.000 -0.0958 0.01461 0.00848 -0.0063 0.4637 0.0201
-2.750 -0.0687 0.01412 0.00789 -0.0060 0.4615 0.0204
-2.500 -0.0415 0.01357 0.00726 -0.0058 0.4591 0.0207
-2.250 -0.0141 0.01305 0.00664 -0.0055 0.4564 0.0209
-2.000 0.0133 0.01254 0.00602 -0.0053 0.4536 0.0213
-1.500 0.0687 0.01173 0.00506 -0.0049 0.4490 0.0222
-1.250 0.0964 0.01129 0.00456 -0.0047 0.4468 0.0226
-1.000 0.1242 0.01093 0.00414 -0.0046 0.4442 0.0230
-0.750 0.1520 0.01064 0.00380 -0.0044 0.4413 0.0234
-0.500 0.1797 0.01040 0.00351 -0.0043 0.4386 0.0238
-0.250 0.2075 0.01019 0.00326 -0.0041 0.4359 0.0240
0.000 0.2353 0.00997 0.00303 -0.0040 0.4337 0.0242
0.250 0.2630 0.00978 0.00283 -0.0039 0.4310 0.0243
0.500 0.2907 0.00961 0.00264 -0.0037 0.4281 0.0244
0.750 0.3183 0.00944 0.00246 -0.0036 0.4251 0.0245
1.000 0.3455 0.00925 0.00224 -0.0034 0.4215 0.0248
1.250 0.3725 0.00898 0.00197 -0.0031 0.4179 0.0254
1.500 0.4000 0.00884 0.00182 -0.0030 0.4127 0.0256
1.750 0.4274 0.00875 0.00169 -0.0028 0.4078 0.0261
2.000 0.4551 0.00867 0.00162 -0.0027 0.4046 0.0266
2.250 0.4829 0.00861 0.00156 -0.0027 0.4009 0.0274
2.500 0.5106 0.00858 0.00153 -0.0026 0.3968 0.0283
2.750 0.5384 0.00858 0.00152 -0.0026 0.3928 0.0290
3.000 0.5663 0.00857 0.00152 -0.0026 0.3892 0.0298
3.250 0.5941 0.00857 0.00153 -0.0026 0.3848 0.0304
3.500 0.6219 0.00860 0.00155 -0.0026 0.3804 0.0310
3.750 0.6497 0.00863 0.00158 -0.0026 0.3763 0.0316
4.000 0.6776 0.00866 0.00162 -0.0026 0.3718 0.0333
4.250 0.7053 0.00871 0.00168 -0.0027 0.3657 0.0373
4.500 0.7271 0.00808 0.00190 -0.0018 0.3596 0.4727
5.000 0.8142 0.00722 0.00229 -0.0089 0.3431 0.9878
5.250 0.8513 0.00738 0.00242 -0.0110 0.3355 0.9905
5.500 0.8868 0.00757 0.00257 -0.0129 0.3238 0.9935
5.750 0.9345 0.00788 0.00277 -0.0176 0.3007 0.9972
6.000 0.9651 0.00823 0.00302 -0.0186 0.2774 0.9981
6.250 0.9932 0.00858 0.00327 -0.0192 0.2572 0.9986
6.500 1.0211 0.00902 0.00360 -0.0197 0.2343 0.9991
6.750 1.0493 0.00942 0.00391 -0.0203 0.2154 0.9996
7.000 1.0760 0.01007 0.00440 -0.0209 0.1833 1.0000
7.250 1.0986 0.01071 0.00488 -0.0206 0.1555 1.0000
7.500 1.1144 0.01250 0.00621 -0.0202 0.0749 1.0000
7.750 1.1306 0.01395 0.00741 -0.0195 0.0181 1.0000
8.000 1.1523 0.01441 0.00788 -0.0189 0.0148 1.0000
8.250 1.1737 0.01487 0.00837 -0.0183 0.0124 1.0000
8.500 1.1950 0.01530 0.00883 -0.0177 0.0116 1.0000
8.750 1.2155 0.01579 0.00935 -0.0170 0.0105 1.0000
9.000 1.2348 0.01637 0.00996 -0.0163 0.0094 1.0000
9.250 1.2531 0.01700 0.01063 -0.0155 0.0085 1.0000
9.500 1.2713 0.01760 0.01127 -0.0147 0.0081 1.0000
9.750 1.2883 0.01827 0.01198 -0.0139 0.0076 1.0000
10.000 1.3035 0.01905 0.01281 -0.0130 0.0072 1.0000
10.250 1.3164 0.02005 0.01387 -0.0123 0.0068 1.0000
10.500 1.3223 0.02170 0.01558 -0.0119 0.0064 1.0000
10.750 1.3178 0.02368 0.01763 -0.0103 0.0061 1.0000
11.000 1.3209 0.02553 0.01955 -0.0096 0.0060 1.0000
11.250 1.3274 0.02723 0.02130 -0.0092 0.0057 1.0000
11.500 1.3323 0.02914 0.02328 -0.0090 0.0055 1.0000
11.750 1.3351 0.03134 0.02554 -0.0088 0.0054 1.0000
12.000 1.3389 0.03346 0.02772 -0.0086 0.0051 1.0000
12.250 1.3410 0.03579 0.03013 -0.0086 0.0050 1.0000
12.500 1.3415 0.03831 0.03270 -0.0086 0.0048 1.0000
12.750 1.3412 0.04092 0.03538 -0.0086 0.0047 1.0000
13.000 1.3411 0.04354 0.03806 -0.0087 0.0045 1.0000
13.250 1.3416 0.04609 0.04068 -0.0087 0.0045 1.0000
13.500 1.3382 0.04918 0.04383 -0.0090 0.0044 1.0000
13.750 1.3318 0.05274 0.04747 -0.0094 0.0042 1.0000
14.000 1.3279 0.05609 0.05090 -0.0099 0.0040 1.0000
14.250 1.3228 0.05970 0.05460 -0.0105 0.0041 1.0000
14.500 1.3229 0.06267 0.05764 -0.0109 0.0040 1.0000
14.750 1.3196 0.06616 0.06120 -0.0116 0.0039 1.0000
15.000 1.3155 0.06978 0.06491 -0.0123 0.0039 1.0000
15.250 1.3143 0.07308 0.06828 -0.0129 0.0038 1.0000
15.500 1.3085 0.07708 0.07237 -0.0138 0.0037 1.0000
15.750 1.3062 0.08057 0.07593 -0.0146 0.0036 1.0000
16.000 1.3028 0.08433 0.07976 -0.0155 0.0035 1.0000
16.250 1.2963 0.08855 0.08407 -0.0166 0.0035 1.0000
16.500 1.2922 0.09247 0.08806 -0.0176 0.0035 1.0000
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Polar data table (+)
Polar graphs
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