GOE 741 AIRFOIL (goe741-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 741 AIRFOIL (goe741-il) Reynolds number: 100,000 Max Cl/Cd: 38.97 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe741-il-100000-n5.txt Download as CSV file: xf-goe741-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 741 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.3136 0.12578 0.11998 -0.0015 0.6668 0.0794
-10.750 -0.3233 0.12246 0.11669 -0.0056 0.6627 0.0804
-10.500 -0.3242 0.11884 0.11308 -0.0083 0.6579 0.0806
-10.000 -0.3200 0.10567 0.09977 -0.0124 0.6507 0.0576
-9.750 -0.3086 0.10304 0.09712 -0.0126 0.6455 0.0570
-9.500 -0.3024 0.09976 0.09384 -0.0138 0.6403 0.0562
-9.250 -0.3003 0.09604 0.09012 -0.0155 0.6361 0.0554
-9.000 -0.3012 0.09198 0.08604 -0.0175 0.6327 0.0545
-8.750 -0.3068 0.08731 0.08134 -0.0201 0.6300 0.0536
-8.500 -0.3121 0.08284 0.07690 -0.0229 0.6263 0.0536
-8.250 -0.3232 0.07778 0.07187 -0.0266 0.6228 0.0534
-8.000 -0.3458 0.07190 0.06597 -0.0296 0.6198 0.0529
-7.500 -0.4265 0.05016 0.04330 -0.0337 0.6174 0.0515
-7.250 -0.4291 0.04552 0.03818 -0.0323 0.6142 0.0515
-7.000 -0.4216 0.04244 0.03474 -0.0308 0.6110 0.0517
-6.750 -0.4081 0.04011 0.03215 -0.0297 0.6066 0.0520
-6.500 -0.3921 0.03812 0.02992 -0.0287 0.6019 0.0525
-6.250 -0.3750 0.03627 0.02780 -0.0275 0.5979 0.0530
-6.000 -0.3566 0.03459 0.02582 -0.0264 0.5945 0.0538
-5.750 -0.3371 0.03306 0.02396 -0.0252 0.5916 0.0549
-5.500 -0.3162 0.03142 0.02197 -0.0243 0.5873 0.0564
-5.250 -0.2941 0.02977 0.01987 -0.0234 0.5830 0.0577
-5.000 -0.2698 0.02889 0.01894 -0.0228 0.5790 0.0586
-4.750 -0.2452 0.02804 0.01794 -0.0222 0.5755 0.0597
-4.500 -0.2202 0.02718 0.01690 -0.0215 0.5726 0.0610
-4.250 -0.1940 0.02637 0.01594 -0.0211 0.5690 0.0625
-4.000 -0.1673 0.02558 0.01497 -0.0208 0.5649 0.0646
-3.750 -0.1415 0.02517 0.01461 -0.0205 0.5610 0.0668
-3.500 -0.1152 0.02468 0.01401 -0.0200 0.5576 0.0698
-3.250 -0.0887 0.02410 0.01330 -0.0195 0.5547 0.0729
-3.000 -0.0625 0.02371 0.01288 -0.0190 0.5519 0.0760
-2.750 -0.0353 0.02336 0.01248 -0.0189 0.5480 0.0806
-2.500 -0.0090 0.02310 0.01231 -0.0186 0.5442 0.0861
-2.250 0.0175 0.02277 0.01197 -0.0183 0.5407 0.0934
-2.000 0.0439 0.02242 0.01156 -0.0178 0.5376 0.1029
-1.750 0.0702 0.02209 0.01120 -0.0173 0.5349 0.1144
-1.500 0.0964 0.02186 0.01102 -0.0169 0.5322 0.1283
-1.250 0.1226 0.02183 0.01118 -0.0168 0.5283 0.1443
-1.000 0.1489 0.02183 0.01130 -0.0166 0.5245 0.1629
-0.750 0.1755 0.02181 0.01130 -0.0163 0.5210 0.1846
-0.500 0.2025 0.02174 0.01122 -0.0160 0.5181 0.2077
-0.250 0.2298 0.02160 0.01111 -0.0157 0.5156 0.2320
0.000 0.2573 0.02155 0.01110 -0.0155 0.5130 0.2575
0.250 0.2842 0.02169 0.01138 -0.0158 0.5085 0.2832
0.500 0.3113 0.02167 0.01149 -0.0158 0.5045 0.3141
0.750 0.3380 0.02147 0.01148 -0.0157 0.5012 0.3663
1.250 0.5557 0.02059 0.01184 -0.0447 0.4929 0.9982
1.500 0.5848 0.02096 0.01221 -0.0454 0.4881 1.0000
1.750 0.6088 0.02121 0.01239 -0.0449 0.4841 1.0000
2.000 0.6330 0.02133 0.01241 -0.0442 0.4807 1.0000
2.250 0.6574 0.02136 0.01229 -0.0433 0.4778 1.0000
2.500 0.6810 0.02164 0.01254 -0.0428 0.4737 1.0000
2.750 0.7037 0.02208 0.01301 -0.0424 0.4681 1.0000
3.000 0.7271 0.02226 0.01314 -0.0417 0.4639 1.0000
3.250 0.7513 0.02229 0.01307 -0.0408 0.4606 1.0000
3.500 0.7759 0.02228 0.01292 -0.0399 0.4580 1.0000
3.750 0.7960 0.02300 0.01379 -0.0395 0.4519 1.0000
4.000 0.8180 0.02335 0.01414 -0.0388 0.4474 1.0000
4.250 0.8412 0.02347 0.01420 -0.0379 0.4439 1.0000
4.500 0.8654 0.02348 0.01413 -0.0370 0.4412 1.0000
4.750 0.8852 0.02408 0.01479 -0.0362 0.4366 1.0000
5.000 0.9039 0.02470 0.01548 -0.0354 0.4314 1.0000
5.250 0.9255 0.02491 0.01569 -0.0344 0.4275 1.0000
5.500 0.9493 0.02490 0.01560 -0.0335 0.4244 1.0000
5.750 0.9681 0.02538 0.01613 -0.0324 0.4199 1.0000
6.000 0.9828 0.02617 0.01702 -0.0312 0.4140 1.0000
6.250 1.0032 0.02638 0.01723 -0.0300 0.4101 1.0000
6.500 1.0268 0.02635 0.01713 -0.0290 0.4071 1.0000
6.750 1.0388 0.02723 0.01811 -0.0275 0.4019 1.0000
7.000 1.0496 0.02808 0.01905 -0.0257 0.3965 1.0000
7.250 1.0691 0.02826 0.01922 -0.0244 0.3928 1.0000
7.500 1.0937 0.02812 0.01902 -0.0235 0.3899 1.0000
7.750 1.0880 0.02984 0.02090 -0.0205 0.3830 1.0000
8.000 1.0977 0.03049 0.02158 -0.0184 0.3782 1.0000
8.250 1.1207 0.03037 0.02142 -0.0173 0.3748 1.0000
8.500 1.1130 0.03195 0.02309 -0.0139 0.3692 1.0000
8.750 1.0792 0.03462 0.02582 -0.0089 0.3626 1.0000
9.000 1.1011 0.03455 0.02571 -0.0076 0.3595 1.0000
9.250 1.1014 0.03588 0.02706 -0.0055 0.3551 1.0000
9.500 1.0527 0.04128 0.03255 -0.0028 0.3454 1.0000
9.750 1.0814 0.04064 0.03190 -0.0018 0.3433 1.0000
10.000 1.1047 0.04051 0.03174 -0.0008 0.3406 1.0000
10.500 1.0354 0.05129 0.04266 0.0011 0.3224 1.0000
10.750 1.1127 0.04557 0.03687 0.0025 0.3257 1.0000
11.000 0.9829 0.06231 0.05378 0.0012 0.3054 1.0000
11.250 1.0692 0.05450 0.04591 0.0035 0.3110 1.0000
11.750 1.0387 0.06280 0.05430 0.0037 0.2959 1.0000
12.000 1.0779 0.06047 0.05195 0.0050 0.2949 1.0000
12.500 1.0390 0.07039 0.06198 0.0043 0.2798 1.0000
12.750 1.0695 0.06910 0.06071 0.0055 0.2789 1.0000
13.250 1.0187 0.08176 0.07351 0.0034 0.2636 1.0000
13.500 1.0464 0.08072 0.07250 0.0044 0.2628 1.0000
14.000 0.9311 0.10494 0.09691 -0.0023 0.2357 1.0000
14.250 0.9480 0.10534 0.09736 -0.0018 0.2337 1.0000
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Polar data table (+)
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