GOE 10K AIRFOIL (goe10k-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 10K AIRFOIL (goe10k-il) Reynolds number: 500,000 Max Cl/Cd: 88.5 at α=1.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe10k-il-500000.txt Download as CSV file: xf-goe10k-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 10K AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.5344 0.08801 0.08578 -0.0101 1.0000 0.0074
-7.500 -0.5382 0.08521 0.08302 -0.0104 1.0000 0.0073
-7.250 -0.5382 0.08181 0.07965 -0.0119 1.0000 0.0077
-7.000 -0.5346 0.07814 0.07599 -0.0143 1.0000 0.0076
-6.750 -0.5294 0.07427 0.07212 -0.0169 1.0000 0.0079
-6.500 -0.5221 0.07027 0.06811 -0.0195 1.0000 0.0081
-6.250 -0.5127 0.06610 0.06391 -0.0220 1.0000 0.0084
-6.000 -0.4989 0.06135 0.05908 -0.0251 1.0000 0.0089
-5.750 -0.4822 0.05669 0.05433 -0.0275 1.0000 0.0091
-5.500 -0.4645 0.05218 0.04968 -0.0289 1.0000 0.0092
-5.250 -0.4499 0.04728 0.04464 -0.0294 1.0000 0.0093
-5.000 -0.4426 0.04402 0.04131 -0.0290 1.0000 0.0097
-4.750 -0.4286 0.04106 0.03824 -0.0285 1.0000 0.0104
-4.500 -0.4110 0.03755 0.03454 -0.0280 1.0000 0.0112
-4.250 -0.3855 0.03389 0.03056 -0.0266 1.0000 0.0126
-4.000 -0.3696 0.02905 0.02538 -0.0256 1.0000 0.0132
-3.750 -0.3534 0.02699 0.02322 -0.0248 1.0000 0.0143
-3.500 -0.3312 0.02522 0.02121 -0.0233 1.0000 0.0168
-3.250 -0.3122 0.02189 0.01752 -0.0224 1.0000 0.0210
-3.000 -0.2864 0.01991 0.01525 -0.0219 0.9995 0.0248
-2.750 -0.2560 0.01757 0.01250 -0.0222 0.9982 0.0309
-2.500 -0.2245 0.01386 0.00806 -0.0198 0.9983 0.0129
-2.250 -0.1968 0.01153 0.00549 -0.0192 0.9977 0.0113
-2.000 -0.1664 0.01047 0.00426 -0.0194 0.9964 0.0118
-1.750 -0.1356 0.00974 0.00345 -0.0200 0.9950 0.0157
-1.500 -0.1050 0.00910 0.00270 -0.0204 0.9928 0.0298
-1.250 -0.0743 0.00863 0.00257 -0.0213 0.9906 0.1213
-1.000 -0.0504 0.00720 0.00265 -0.0215 0.9887 0.5696
-0.750 0.0013 0.00611 0.00266 -0.0267 0.9944 1.0000
-0.500 0.0379 0.00615 0.00260 -0.0288 0.9912 1.0000
-0.250 0.0826 0.00620 0.00257 -0.0327 0.9878 1.0000
0.000 0.1217 0.00616 0.00248 -0.0352 0.9833 1.0000
0.250 0.1668 0.00609 0.00239 -0.0391 0.9791 1.0000
0.500 0.2055 0.00599 0.00230 -0.0415 0.9738 1.0000
0.750 0.2458 0.00589 0.00221 -0.0443 0.9691 1.0000
1.000 0.2901 0.00561 0.00198 -0.0477 0.9598 1.0000
1.250 0.3344 0.00528 0.00173 -0.0510 0.9467 1.0000
1.500 0.3818 0.00503 0.00157 -0.0551 0.9284 1.0000
1.750 0.4416 0.00499 0.00127 -0.0614 0.8051 1.0000
2.000 0.4352 0.00698 0.00150 -0.0539 0.3939 1.0000
2.250 0.4461 0.00850 0.00188 -0.0511 0.1006 1.0000
2.500 0.4656 0.00939 0.00244 -0.0492 0.0182 1.0000
2.750 0.4869 0.01009 0.00325 -0.0476 0.0129 1.0000
3.000 0.5083 0.01084 0.00409 -0.0459 0.0115 1.0000
3.250 0.5286 0.01192 0.00526 -0.0440 0.0111 1.0000
3.500 0.5489 0.01359 0.00696 -0.0422 0.0099 1.0000
3.750 0.5731 0.01511 0.00867 -0.0407 0.0101 1.0000
4.500 0.6341 0.02760 0.02260 -0.0329 0.0189 1.0000
4.750 0.6619 0.02708 0.02230 -0.0313 0.0159 1.0000
5.000 0.6800 0.02916 0.02458 -0.0296 0.0145 1.0000
5.250 0.6877 0.03444 0.03005 -0.0276 0.0134 1.0000
5.500 0.7114 0.03632 0.03245 -0.0244 0.0122 1.0000
5.750 0.7271 0.03963 0.03603 -0.0223 0.0111 1.0000
6.000 0.7398 0.04264 0.03922 -0.0207 0.0104 1.0000
6.500 0.7111 0.03835 0.03551 -0.0149 0.0098 1.0000
6.750 0.7173 0.04314 0.04053 -0.0132 0.0097 1.0000
7.000 0.7234 0.04839 0.04601 -0.0121 0.0095 1.0000
7.250 0.7248 0.05388 0.05166 -0.0117 0.0093 1.0000
7.500 0.7210 0.05906 0.05694 -0.0118 0.0091 1.0000
7.750 0.7120 0.06404 0.06199 -0.0121 0.0088 1.0000
8.000 0.6971 0.06822 0.06620 -0.0122 0.0090 1.0000
8.250 0.6827 0.07269 0.07069 -0.0141 0.0094 1.0000
8.500 0.6710 0.07800 0.07599 -0.0176 0.0095 1.0000
8.750 0.6633 0.08336 0.08134 -0.0202 0.0095 1.0000
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