GOE 673 AIRFOIL (goe673-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: GOE 673 AIRFOIL (goe673-il) Reynolds number: 100,000 Max Cl/Cd: 55.53 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe673-il-100000-n5.txt Download as CSV file: xf-goe673-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 673 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4214 0.09348 0.08866 -0.0416 1.0000 0.0247
-9.000 -0.4252 0.08923 0.08449 -0.0432 1.0000 0.0243
-8.750 -0.4318 0.08495 0.08030 -0.0446 1.0000 0.0237
-8.500 -0.4423 0.08073 0.07619 -0.0456 1.0000 0.0231
-8.250 -0.4778 0.07018 0.06577 -0.0520 1.0000 0.0195
-8.000 -0.4925 0.06627 0.06177 -0.0526 1.0000 0.0190
-7.750 -0.5040 0.06324 0.05863 -0.0518 1.0000 0.0188
-7.500 -0.5127 0.06010 0.05535 -0.0506 1.0000 0.0187
-7.250 -0.5191 0.05656 0.05163 -0.0496 1.0000 0.0188
-7.000 -0.5219 0.05280 0.04764 -0.0486 1.0000 0.0188
-6.750 -0.5071 0.04716 0.04155 -0.0517 0.9961 0.0191
-6.500 -0.4842 0.04212 0.03599 -0.0548 0.9913 0.0194
-6.250 -0.4558 0.03804 0.03131 -0.0573 0.9872 0.0193
-6.000 -0.4278 0.03453 0.02724 -0.0589 0.9826 0.0193
-5.750 -0.3970 0.03130 0.02348 -0.0607 0.9789 0.0195
-5.500 -0.3660 0.02858 0.02034 -0.0620 0.9751 0.0200
-5.250 -0.3352 0.02647 0.01793 -0.0632 0.9708 0.0213
-5.000 -0.3011 0.02485 0.01607 -0.0648 0.9674 0.0240
-4.750 -0.2697 0.02333 0.01429 -0.0656 0.9630 0.0257
-4.500 -0.2379 0.02201 0.01280 -0.0665 0.9583 0.0273
-4.250 -0.2029 0.02085 0.01146 -0.0683 0.9548 0.0309
-4.000 -0.1718 0.01987 0.01034 -0.0693 0.9495 0.0362
-3.750 -0.1386 0.01895 0.00928 -0.0706 0.9445 0.0458
-3.500 -0.1033 0.01735 0.00834 -0.0729 0.9412 0.1773
-3.250 -0.0740 0.01622 0.00817 -0.0742 0.9360 0.3849
-3.000 -0.0440 0.01567 0.00806 -0.0748 0.9303 0.5136
-2.750 -0.0089 0.01537 0.00774 -0.0761 0.9263 0.5648
-2.500 0.0175 0.01519 0.00760 -0.0756 0.9182 0.6133
-2.250 0.0510 0.01498 0.00737 -0.0765 0.9134 0.6662
-2.000 0.0773 0.01484 0.00722 -0.0759 0.9052 0.7122
-1.750 0.1086 0.01463 0.00701 -0.0761 0.8996 0.7580
-1.500 0.1330 0.01447 0.00691 -0.0748 0.8911 0.8092
-1.250 0.1646 0.01425 0.00673 -0.0746 0.8855 0.8732
-1.000 0.2019 0.01414 0.00659 -0.0762 0.8777 0.9392
-0.750 0.2498 0.01396 0.00629 -0.0803 0.8721 0.9875
-0.500 0.2802 0.01391 0.00610 -0.0811 0.8623 1.0000
-0.250 0.3094 0.01385 0.00592 -0.0815 0.8536 1.0000
0.000 0.3386 0.01379 0.00576 -0.0818 0.8446 1.0000
0.250 0.3651 0.01380 0.00569 -0.0816 0.8339 1.0000
0.500 0.3929 0.01379 0.00561 -0.0817 0.8238 1.0000
0.750 0.4224 0.01374 0.00550 -0.0819 0.8147 1.0000
1.000 0.4495 0.01376 0.00547 -0.0818 0.8034 1.0000
1.250 0.4764 0.01379 0.00547 -0.0816 0.7917 1.0000
1.500 0.5039 0.01382 0.00549 -0.0815 0.7801 1.0000
1.750 0.5317 0.01385 0.00550 -0.0814 0.7685 1.0000
2.000 0.5597 0.01387 0.00551 -0.0814 0.7567 1.0000
2.250 0.5876 0.01391 0.00557 -0.0813 0.7445 1.0000
2.500 0.6145 0.01400 0.00566 -0.0811 0.7310 1.0000
2.750 0.6415 0.01409 0.00577 -0.0808 0.7171 1.0000
3.000 0.6685 0.01419 0.00590 -0.0806 0.7029 1.0000
3.250 0.6955 0.01431 0.00608 -0.0803 0.6883 1.0000
3.500 0.7223 0.01445 0.00625 -0.0800 0.6734 1.0000
3.750 0.7486 0.01462 0.00647 -0.0797 0.6577 1.0000
4.000 0.7748 0.01482 0.00676 -0.0793 0.6417 1.0000
4.250 0.8008 0.01504 0.00707 -0.0789 0.6258 1.0000
4.500 0.8268 0.01528 0.00741 -0.0785 0.6101 1.0000
4.750 0.8526 0.01554 0.00780 -0.0780 0.5943 1.0000
5.000 0.8776 0.01581 0.00820 -0.0774 0.5746 1.0000
5.250 0.8946 0.01611 0.00831 -0.0748 0.5148 1.0000
5.500 0.9064 0.01684 0.00854 -0.0717 0.4176 1.0000
5.750 0.9121 0.01841 0.00923 -0.0681 0.2675 1.0000
6.000 0.9058 0.02181 0.01108 -0.0636 0.0508 1.0000
6.250 0.9208 0.02317 0.01244 -0.0616 0.0356 1.0000
6.500 0.9362 0.02438 0.01380 -0.0596 0.0299 1.0000
6.750 0.9503 0.02558 0.01521 -0.0577 0.0256 1.0000
7.000 0.9588 0.02715 0.01692 -0.0549 0.0239 1.0000
7.250 0.9692 0.02851 0.01847 -0.0522 0.0229 1.0000
7.500 0.9791 0.02993 0.02004 -0.0496 0.0217 1.0000
7.750 0.9901 0.03135 0.02159 -0.0471 0.0199 1.0000
8.000 1.0043 0.03295 0.02328 -0.0452 0.0184 1.0000
8.250 1.0246 0.03491 0.02533 -0.0441 0.0175 1.0000
8.500 1.0514 0.03724 0.02781 -0.0439 0.0168 1.0000
8.750 1.0777 0.03984 0.03061 -0.0436 0.0163 1.0000
9.000 1.0980 0.04248 0.03348 -0.0427 0.0156 1.0000
9.250 1.1129 0.04572 0.03696 -0.0416 0.0146 1.0000
9.500 1.1225 0.05042 0.04204 -0.0400 0.0139 1.0000
9.750 1.1251 0.05375 0.04572 -0.0373 0.0136 1.0000
10.000 1.1248 0.05667 0.04907 -0.0343 0.0136 1.0000
10.250 1.1198 0.05968 0.05238 -0.0311 0.0135 1.0000
10.500 1.1118 0.06285 0.05584 -0.0280 0.0136 1.0000
10.750 1.1015 0.06624 0.05950 -0.0253 0.0136 1.0000
11.000 1.0891 0.06987 0.06338 -0.0231 0.0136 1.0000
11.250 1.0749 0.07385 0.06759 -0.0215 0.0137 1.0000
11.750 1.0461 0.08209 0.07632 -0.0204 0.0138 1.0000
12.000 1.0314 0.08641 0.08083 -0.0208 0.0138 1.0000
12.250 1.0163 0.09112 0.08574 -0.0221 0.0139 1.0000
12.500 1.0011 0.09626 0.09105 -0.0241 0.0140 1.0000
12.750 0.9859 0.10185 0.09682 -0.0268 0.0141 1.0000
13.000 0.9707 0.10801 0.10313 -0.0304 0.0142 1.0000
13.250 0.9555 0.11489 0.11016 -0.0348 0.0143 1.0000
13.500 0.9403 0.12282 0.11823 -0.0402 0.0145 1.0000
13.750 0.9245 0.13253 0.12806 -0.0470 0.0148 1.0000
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Polar data table (+)
Polar graphs
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